Process and machine for reducing a drag component of a horizontal stabilizer on an aircraft

ABSTRACT

A process and a machine for improving a performance of a particular model of an aircraft, via reducing a size of a horizontal stabilizer for the particular model of the aircraft, the process comprising augmenting a nose-down moment, for the particular aircraft model, provided by a reduced horizontal stabilizer for the particular aircraft model, via addition of an ailevatoron mixer.

CROSS REFERENCE TO PARENT APPLICATION

This application is a continuation-in-part (CIP) of and claims priorityto the following U.S. patent application entitled “Process and Machinefor Reducing a Drag Component of a Horizontal Stabilizer on anAircraft,” Ser. No. 15/279,414, attorney docket no. 16-0668-US-NP, filedSep. 28, 2016, which is incorporated herein by reference.

BACKGROUND INFORMATION 1. Field

The present disclosure relates generally to aircraft and, in particular,to controlling movement of the aircraft. Still more particularly, thepresent disclosure relates to reducing a drag component of a horizontalstabilizer of an aircraft via a process and machine for producing apitch moment for an aircraft.

2. Background

Aircraft designers and operators face a constant challenge to create aprocess, machine, manufacture, and/or composition of matter that producea technical effect that will reduce a cost and/or improve a performance,of an aircraft. Without limitation, reference to the aircraft mayindicate not only the aircraft as a whole, but may also indicate aparticular part and/or parts of the aircraft. Without limitation,reducing a cost can include reducing a cost to manufacture, maintain,and/or operate the aircraft. Without limitation, improving theperformance of the aircraft can include an improvement inresponsiveness, efficiency, and/or an increase in reliability for theaircraft. Without limitation, improving the efficiency of the aircraftmay include reducing a weight and/or drag of the aircraft, and/orimproving a fuel efficiency of the aircraft. Thus, improving a process,machine, manufacture, and/or composition of matter for producing a pitchmoment for an aircraft provides the technical effect of improvingresponsiveness, efficiency, and/or reliability for the aircraft.

While various structures of an aircraft generate aerodynamic forces thataffect pitch control of the aircraft, most commonly, fixed wing aircraftcontrol a nose-up or a nose-down pitch moment by generating anaerodynamic force on a horizontal stabilizer mounted at a tail of theaircraft. When the aircraft is in flight the pitch moments are evaluatedas acting about a center of gravity of the aircraft. Pitch moments arealso typically evaluated about a center of gravity for the aircraft whenthe aircraft is on the ground. However, to illustrate their effect onaircraft takeoff rotation, when the aircraft has the wheels of its maingear supporting the aircraft on a surface, the pitch moment may beevaluated as acting about an axis where the main gear of the aircraftcontact the surface.

For a tricycle wheeled aircraft to stand or roll on a surface withoutsuffering a tail portion of the aircraft dropping toward the surface, atleast far enough to lift a nosewheel of the aircraft off of the surface,the center of gravity of the aircraft must be located forward (towardthe aircraft nose) of the axis where the main gear of the aircraftcontact the surface. Without limitation, the surface may be a runway,taxiway, and/or a ramp area.

For a tricycle wheeled aircraft to rotate its nosewheel up off thesurface at takeoff (commonly referred to as rotation), the aircraft mustbe able to provide a moment, about the axis where the aircraft's maingear contact the runway, that is sufficient to lift the nosewheel of theaircraft up off the runway (referred to herein as a rotation moment), toa specified distance that sets a specified takeoff attitude for theaircraft relative to the runway and provides a specified takeoff pitchand/or takeoff angle of attack in the relative wind for the wings of theaircraft. Traditionally, the rotation moment that lifts the aircraftnosewheel off the runway has been provided by an aerodynamic forcegenerated by a horizontal stabilizer at the tail of the aircraft.

Regulations, from various agencies around the globe, for certificationof aircraft, particularly for aircraft used for commercial transport,require that the aircraft design must be able to lift the nose of theaircraft up to a position that results in liftoff from the runway at aspecified airspeed, called rotation speed, also represented as V_(r). Anon-limiting example of such regulations in the United States may befound in the Federal Aviation Regulations (FARs), such as withoutlimitation, FAR Part 25, and Federal Aviation Administration AdvisoryCirculars issued thereto. In particular, without limitation,requirements driven by 14 CFR §§ 25.101-115 may control rotation speedand other performance requirements that may affect without limitation,at least a size, a shape, and a location of a flight control surface onan aircraft, and systems that control and/or actuate the flight controlsurface on the aircraft.

More specifically, traditionally nearly all the rotation moment isgenerated by a horizontal stabilizer of the aircraft via producing anaerodynamic force from the horizontal stabilizer. There is a particularset of circumstances where the rotation moment required will be at amaximum value. As a non-limiting example, for a given rotation speed,when a center of gravity of the aircraft is located at its forward mostallowed point, and the aircraft weight is at a maximum allowed value fortakeoff, the rotation moment required that is sufficient to lift thenosewheel of the aircraft up off the runway to a specified distance thatsets a specified takeoff attitude for the aircraft, may be at a maximumvalue. It is understood by one of skill in the art that the givenrotation speed may be affected at least by, without limitation, aconfiguration of the aircraft, a thrust of the aircraft, rollingfriction, and atmospheric conditions such as without limitation,altitude and temperature.

Moreover, for any particular aircraft model, some combination of:allowable forward (toward the nose of the aircraft from the axis wherethe main gear contact the runway) center of gravity, aircraft grossweight, and designated rotation speed will result in a value of rotationmoment that must be generated to lift the nose of the aircraft up to aposition that results in a takeoff attitude for the aircraft that at thedesignated rotation speed results in a liftoff of the aircraft from therunway, that is at a maximum value for the particular model aircraft.

Factors that determine the aerodynamic force that the horizontalstabilizer can produce to generate the rotation moment include anairspeed of the aircraft, a size of the horizontal stabilizer, a shapeof the horizontal stabilizer, and an angle of attack of the horizontalstabilizer (in the relative wind impacting a leading edge of thehorizontal stabilizer). Generally, when the nose gear and the main gearare in contact with the surface the aircraft is taking off from, theangle of attack of the horizontal stabilizer is set by a stabilizer trimposition. Changes in the shape of the horizontal stabilizer (inoperation, not in design phase) are generally limited to a deflection ofa moveable flight control panel located in the horizontal stabilizer,known as an elevator.

In flight, for a horizontal stabilizer, of a particular airfoil shapeset at a particular trim position and at a particular angle of attack tothe relative wind meeting the horizontal stabilizer, increasing theairspeed will increase an aerodynamic force produced by the horizontalstabilizer, and thus increase a moment the horizontal stabilizer canproduce, commonly measured about the center of gravity of the aircraft.

Thus, for a horizontal stabilizer of a particular shape at a given trimposition using an elevator deflection that produces a maximumaerodynamic force that the horizontal stabilizer is capable of producingat the rotation speed, the maximum value of aircraft nose-up moment thatthe horizontal stabilizer can generate is proportional to the size ofthe horizontal stabilizer. Hence, for an aircraft where the momentrequired for rotation is provided principally by an aircraft nose-uppitch moment generated by the aerodynamic force generated by thehorizontal stabilizer, as is the case for nearly all existing fixed wingtricycle gear configured aircraft, and particularly so for existingcommercial transport aircraft, a limit to the smallest size currentlyallowable for a horizontal stabilizer may be set by a required maximumrotation moment at a set rotation speed for a combination of aparticular takeoff weight for the aircraft with a particular forwardlocation of the center of gravity of the aircraft.

Additionally, as one of ordinary skill in the art appreciates,increasing a size of the horizontal stabilizer increases the profiledrag produced by the horizontal stabilizer. Further, for a given sizedhorizontal stabilizer, profile drag produced by the horizontalstabilizer increases as airspeed of the aircraft is increased. Thus, forany given sized horizontal stabilizer, the profile drag produced by thehorizontal stabilizer will increase as the aircraft airspeed increases.

Once the aircraft is off the runway, all moments that raise or lower anaircraft nose relative to the horizon, or pitch moments, are normallycalculated as acting, not about any point on the main gear, but instead,about the center of gravity of the aircraft. At a given airspeed andwith the horizontal stabilizer fixed at a given distance from the centerof gravity, either a larger angle, relative to a relative wind at thehorizontal stabilizer, a greater camber of the horizontal stabilizer, ora larger size of the horizontal stabilizer is generally needed toincrease the moment generated about the horizontal stabilizer.Increasing an effective camber of the horizontal stabilizer may becaused by an increased deflection of the elevator on the horizontalstabilizer. In particular, during takeoff roll for an aircraft, a size,a shape, and an angle of attack for a given horizontal stabilizerdetermine how much pitch change (rotation) the aircraft can achieve atany given airspeed.

Traditionally, and particularly so for commercial transport aircraft,once off the runway, pitch moments approaching a magnitude required forthe rotation moment are not required. Even if such a magnitude wasrequired in flight, at least because the airspeed of the aircraft inflight is greater than rotation speed, such moments could be provided bya much smaller sized horizontal stabilizer than is needed at rotationspeed.

As the airspeed of the aircraft increases above rotation speed, the sizeneeded for the horizontal stabilizer to produce a particular value ofaerodynamic force decreases, because a force produced by a particularsize horizontal stabilizer at a given angle of attack increases withincreased speed. Hence, at speeds higher than rotation speed, ahorizontal stabilizer of a fixed size is able to generate greateraerodynamic forces and thus greater moments about a center of gravity ofthe aircraft than the horizontal stabilizer was capable of at rotationspeed. Thus, for aircraft designs that lack the technical benefits ofthe embodiments described herein, after rotation at takeoff, theaircraft pays drag penalties because the horizontal stabilizer isoversized for all the remaining performance requirements for the flightof the aircraft.

Thus, if the full size of a particular horizontal stabilizer is notneeded to produce the magnitudes of moments needed for aircraftperformance requirements throughout the rest of the flight afterrotation, using a particular size horizontal stabilizer to generate theaerodynamic force that generates the required rotation moment at a setrotation speed, results in flying the entire remainder of the flightafter rotation with a horizontal stabilizer that is much larger thanneeded to produce the aerodynamic forces that will be required from theparticular horizontal stabilizer for the rest of the flight. In otherwords, without incorporating an embodiment of a novel machine and/orprocess described herein, an aircraft, after rotation, may carry a dragload from a horizontal stabilizer that is larger than would be requiredfor those aircraft by an embodiment of the novel machine and/or processdescribed herein.

Accordingly, this oversized (traditionally sized) horizontal stabilizer,creates the technical deficiency of producing drag penalties thatnegatively impact the aircraft performance and fuel economy of theaircraft up until, and then after, rotation. Further, these dragpenalties from the traditional horizontal stabilizer are not offset byany improvements to performance of the aircraft for the expected flightenvelope for the aircraft after rotation.

The oversized horizontal stabilizer also weighs more than a novelhorizontal stabilizer of a reduced size would. Increased weight of anaircraft tends to increase the aircraft's fuel consumption. Hence, foraircraft with a traditional horizontal stabilizer, drag, aircraftweight, and resulting fuel economy for the aircraft all exist intechnical deficiency as compared to an aircraft that might be able toproduce the required maximum rotation moment with a horizontalstabilizer that is smaller than the traditional horizontal stabilizer(referred to hereinafter as a reduced horizontal stabilizer). In otherwords, large numbers of aircraft flying today, after rotation, carry anexcessive amount of unproductive weight from a horizontal stabilizerthat is heavier than would be required for those aircraft by embodimentsof the novel machine and process described herein. In other words,without incorporating an embodiment of a novel machine and/or processdescribed herein, an aircraft, after rotation, may an excessive amountof unproductive weight from a horizontal stabilizer that is larger thanwould be required for those aircraft by an embodiment of the novelmachine and/or process described herein.

Therefore, it would be desirable to have a process and machine that takeinto account at least some of the issues discussed above, as well asother possible issues. For example, it would be desirable to have aprocess and machine that overcome the technical deficiencies resultingfrom a current size requirement for a horizontal stabilizer determinedby current technologies for generating a rotation moment for an aircraftduring takeoff.

SUMMARY

Embodiments herein describe at least a process for reducing a size of ahorizontal stabilizer for a particular aircraft model, the processincluding augmenting a nose-up moment, for the aircraft, provided by ahorizontal stabilizer of a reduced size for the particular aircraftmodel. Another embodiment includes a process for improving the fuelefficiency for a given aircraft via reducing a size of a horizontalstabilizer for a particular aircraft model. Another embodiment includesa machine configured to reduce a size of a horizontal stabilizer for aparticular aircraft model while sustaining a nose-up moment, requiredfor takeoff, for an aircraft of a particular model loaded at maximumgross takeoff weight with a center of gravity located at a maximumforward location allowed, such that the machine comprises an aileroncontrol system configured to symmetrically deflect ailevatorons, locatedaft of an axis of contact of a main gear of the aircraft, away from arunway.

Additional embodiments may show a process for supplementing a pitchmoment generated by a given horizontal stabilizer for a given aircraft.Further embodiments may provide a process of reducing a takeoff/rotationairspeed for an aircraft via a machine for supplementing a pitch momentproduced by a given horizontal stabilizer for a given aircraft.

BRIEF DESCRIPTION OF THE DRAWINGS

Features believed novel and/or characteristic of the illustrativeembodiments are recited in the appended claims. Understanding of theillustrative embodiments, as well as a preferred mode of use, furtherobjectives and features thereof, will be enhanced by reference to thefollowing detailed description of illustrative embodiments of thepresent disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is an illustration a side-view of an aircraft with a nosewheeland main gear in contact with a surface, depicted in accordance with anillustrative embodiment;

FIG. 2 is an illustration of a schematic representation of relativemagnitudes of rotation moment control for an aircraft, and componentscontributing thereto, is depicted in accordance with an illustrativeembodiment;

FIG. 3 presents a graph that represents the relative magnitude of anupward deflection of a trailing edge of an ailevatoron relative to amagnitude of a trailing edge of an elevator deflection in a directionthat generates a nose-up pitch for an aircraft, in accordance with anillustrative embodiment;

FIG. 4 is an illustration of a diagram of a data processing system forimplementing an ailevatoron mixer is depicted in accordance with anadvantageous embodiment;

FIG. 5 is a functional schematic diagram of an ailevatoron mixer controland logic, in accordance with an advantageous embodiment;

FIG. 6 is an illustration of a flowchart for operations performed by oneembodiment for reducing a size of a horizontal stabilizer for aparticular aircraft model, in accordance with an advantageousembodiment;

FIG. 7 is an illustration of a flowchart for operations performed by oneembodiment for reducing a size of a horizontal stabilizer for aparticular model of an aircraft, in accordance with an advantageousembodiment;

FIG. 8 is an illustration of a flowchart for operations performed by oneembodiment for increasing the fuel efficiency for a particular aircraftmodel, in accordance with an advantageous embodiment;

FIG. 9 shows a diagram illustrating operations of an embodiment for anaircraft manufacturing and service method, depicted in accordance withan advantageous embodiment;

FIG. 10 shows a plot of an angle of attack versus a maximum nose-downpitching moment coefficient (Cm) for an example configuration of anexample aircraft in-flight, depicted in accordance with an advantageousembodiment;

FIG. 11 shows an aircraft in flight with a wing at a stall angle ofattack, depicted in accordance with an advantageous embodiment;

FIG. 12A shows a plot of an angle of attack versus a maximum nose-downpitching moment coefficient (Cm) for the example configuration of theexample aircraft in-flight, depicted in accordance with an advantageousembodiment;

FIG. 12B shows a plot of an angle of attack versus a maximum nose-downpitching moment coefficient (Cm) for the example configuration of theexample aircraft in-flight, depicted in accordance with an advantageousembodiment;

FIG. 13 shows a plot of an angle of attack versus a maximum nose-downpitching moment coefficient (Cm) for an example aircraft in-flight witha center of gravity located at various positions, depicted in accordancewith an advantageous embodiment;

FIG. 14 shows a flowchart for a process for improving a performance of aparticular model of an aircraft via reducing a size of a horizontalstabilizer on the particular model of the aircraft to form a smallerhorizontal stabilizer, and sustaining a required stall recoverynose-down pitching moment capability of the aircraft, in accordance withan illustrative embodiment; and

FIG. 15 shows a flowchart for a process for expanding an allowable rangefor a center of gravity of a particular model of an aircraft andretaining a configuration of the particular model of the aircraft.

DETAILED DESCRIPTION

The illustrative embodiments recognize and take into account one or moredifferent considerations. Aircraft components may produce lift and drag.For example, the illustrative embodiments recognize and take intoaccount that reducing a value of aerodynamic drag (drag for short) forany component of an aircraft may increase the fuel efficiency of theaircraft. Lift producing components may produce an induced drag due tolift production and a profile drag due to the profile (also known as ashape, a size, or a form of the component). Generally the amount ofprofile drag a component produces increases as the airspeed of theaircraft increases. Therefore, to lower profile drag contributed by acomponent of the aircraft it may be desirable to reduce a size of acomponent, such as without limitation a horizontal stabilizer of theaircraft.

Reducing a size of an aircraft component may also reduce a weight of thecomponent and thus reduce a weight of the aircraft. Reducing the weightof the aircraft may increase the fuel efficiency of the aircraft. Thus,it may be desirable, for any given aircraft to reduce a size and weightof the horizontal stabilizer for that aircraft, and thereby reduce thedrag and increase fuel efficiency of the aircraft.

The illustrative embodiments recognize and take into account thatreducing the size of the horizontal stabilizer for a particular aircraftmodel may reduce an ability of the horizontal stabilizer to producepitch moments about center of gravity and reduce an ability of thehorizontal stabilizer to generate a rotation moment about an axis wherethe main gear contact takeoff surface.

Thus, in order to produce a maximum rotation moment required forrotation and takeoff for a particular aircraft model, a process and/ormachine may be needed to supplement the rotation moment produced by ahorizontal stabilizer (referred to hereinafter as stab moment—for atraditionally sized horizontal stabilizer) of a reduced size (referredto hereinafter as reduced moment produced by a reduced horizontalstabilizer) for the particular aircraft model. As used herein, aparticular aircraft model may refer to a particular aircraft type, suchas without limitation a B777, or to a particular series of a typeaircraft, such as without limitation B777-300, or to a generation of anaircraft design, such as without limitation, the B737 MAX.

The illustrative embodiments recognize and take into account thatdespite the ability of the horizontal stabilizer of any given shape toincrease pitch moments produced by the horizontal stabilizer as anairspeed of the aircraft increases, the maximum pitch moment requirementthat a traditional horizontal stabilizer must produce throughout itsentire flight envelope is often the rotation moment needed at takeoff.Thus, at least as explained above, a size and/or shape (also known asform or profile) as well as the weight of the traditional horizontalstabilizer create technical inefficiencies at least of increased dragand fuel consumption of the aircraft throughout the rest of the flightafter takeoff.

The illustrative embodiments also recognize and take intoconsiderations, that sizing requirements for a horizontal stabilizer maybe impacted at least by: a range allowed for a location for a center ofgravity of the aircraft, and by requirements to control aircraftnose-down pitch in flight. The requirements may be driven by an airspeednear or at stall identification, such as without limitation. UnitedStates Federal Aviation Regulations presented in 14 CFR § 25.145 and/or14 CFR § 25.201 and 25.103. Further, in aircraft where ailerons remaineffective in stall conditions, such as, without limitation, via aleading edge auto-slat deployment system, and/or a location and/orsizing of an aileron and/or ailevatoron, a stall recovery system may beapplied that would reduce a size requirement of a horizontal stabilizerthat is driven by nose-down pitch authority requirements.

Additionally, the illustrative embodiments recognize and take intoaccount that fly-by wire control systems using buses, such as those usedin computers, are becoming more common in aircraft. For example, specialflight control programs in a computer processor may send commands tospecial actuator control programs in processors that control devices inthe aircraft. Actuator control programs may control, for example, aflight control surface, an engine, or some other suitable device in theaircraft that may affect a change in pitch attitude or rate of anaircraft.

The illustrative embodiments recognize and take into account that a busmay be a parallel bus or a serial bus. When a parallel bus is used,units of data, such as a word, may be carried on multiple paths in thebus.

Moreover, the illustrative embodiments provide a method and apparatusfor controlling flight control surfaces on an aircraft. Such methods andapparatus may include a data bus system, and actuator control programsand/or flight control programs specially programmed in processors. Adata bus system may be located in an aircraft. The actuator controlprograms may be in communication with the data bus system.

An actuator control program may control positioning of a group of flightcontrol surfaces on the aircraft using commands on the data bus systemthat are directed to the actuator control program. Control of flightcontrol actuators may be commanded by a flight control system that maycontain flight control programs that may be connected to the data bussystem.

The flight control programs generate and send the commands onto the bussystem to control the flight control surfaces on the aircraft. Thecommands for a flight control surface are directed towards a group ofactuator control programs on processors assigned to the actuators of theflight control surfaces.

A “group of,” as used herein with reference to items, means one or moreitems. For example, a group of actuator control modules is one or moreactuator control modules.

As used herein, the phrase “at least one of,” when used with a list ofitems, means different combinations of one or more of the listed itemsmay be used and only one of each item in the list may be needed. Inother words, at least one of means any combination of items and numberof items may be used from the list but not all of the items in the listare required. The item may be a particular object, thing, or a category.

For example, without limitation, “at least one of item A, item B, oritem C” may include item A, item A and item B, or item B. This examplealso may include item A, item B, and item C or item B and item C. Ofcourse, any combinations of these items may be present. In otherexamples, “at least one of” may be, for example, without limitation, twoof item A; one of item B; and ten of item C; four of item B and seven ofitem C; or other suitable combinations.

With reference now to the figures, and in particular, with reference toFIG. 1, an illustration of a side-view of an aircraft with a nosewheeland main gear in contact with a surface is depicted in accordance withan illustrative embodiment. In this illustrative example, aircraft 100is shown standing on surface 102 with the wheels of (right side shown)main gear 104 contacting surface 102 at axis 106 while nosewheel 108 isalso in contact with surface 102. Axis 106 is formed by a line betweenthe contact point of a left side main gear (aligned and hidden behindright side main gear shown) and the contact point of the right main gear104 with surface 102. Without limitation, surface 102 may be a takeoffsurface, a runway, a taxiway, or a parking apron or other ramp area.

Center of gravity 110 is shown located within range 112 of allowablelocations for center of gravity 110 for aircraft 100. Range 112 oflocations allowable for center of gravity 110 for aircraft 100 may bedetermined by performance considerations for aircraft 100 andregulations. Regulations may include without limitation United StatesFederal Aviation Regulations Part 25.

Aerodynamic force 114 is generated by (traditionally sized) horizontalstabilizer 116 located at tall 118 of aircraft 100. Aerodynamic force114 generated by (traditionally sized) horizontal stabilizer 116 may bea representation of a summation of a lift force and a drag forceproduced by (traditionally sized) horizontal stabilizer 116.

For aircraft 100 as shown in FIG. 1, moment 120, also called weightmoment 120, acts to hold nosewheel 108 on surface 102 by acting aboutaxis 106, caused by weight 122 acting through center of gravity 110.During initial takeoff roll for aircraft 100, weight moment 120 isgreater than stab moment 124 acting about axis 106, caused byaerodynamic force 114 on (traditionally sized) horizontal stabilizer116. Hence, with weight moment 120 greater than stab moment 124,nosewheel 108 remains on surface 102 as aircraft 100 stands or rollsalong surface 102.

At rotation speed, rotation moment 130 must be generated that will liftnosewheel 108 above surface 102 and place aircraft 100 in a takeoffattitude. For any particular aircraft model, at rotation speed,aerodynamic forces generated by other (non-horizontal stabilizer)components of the aircraft, such as without limitation wing 126, maingear 104, a fuselage, and/or thrust, may contribute to moments aboutaxis 106. Wing 126 shown is a wing on a right side of aircraft 100,however wing 126 and components thereon, discussed below in furtherdetail, may be considered as representative for a left wing on aircraft100 as well.

Forces generated by non-horizontal stabilizer components of aircraft 100may affect rotation of aircraft 100 and rotation moment 130. Withoutlimitation, such forces may include at least thrust and lift. Forclarity of emphasis on the technical effects of the machine and/orprocess embodiments described herein, forces generated by othernon-horizontal stabilizer components of aircraft 100 not specificallydetailed within the descriptions below are incorporated into rotationmoment 130 described herein, and thus are not specified in detail in thepresent review of the illustrative embodiments, or included in FIG. 1,with the exception of forces and moments generated and/or affected bythe novel embodiment(s) introduced herein.

Thus, as described above, if (traditionally sized) horizontal stabilizer116 has its size reduced to form reduced horizontal stabilizer 132 thenaerodynamic force 114 will be replaced by aerodynamic force 134, whichis smaller than aerodynamic force 114. As a result stab moment 124 willbe replaced by reduced moment 136. Reduced moment 136 will have a lowermagnitude than stab moment 124, and thus will need a supplementarymoment added to produce the equivalent magnitude required to equal themagnitude of rotation moment 130.

When aircraft 100 has wing 126 in a configuration, such as withoutlimitation a swept configuration, where ailevatoron 138 on wing 126 islocated aft of axis 106, then when ailevatoron 138 produces force 140that acts downward toward surface 102 and aft (toward tail 118 from axis106) of axis 106, ailevatoron 138 generates ailevatoron moment 142.Ailevatoron 138 is an aerodynamic surface on a wing that may replaceand/or supplement a traditional aileron. An existing aileron on wing 126may be commanded to a use that may re-designate the aileron asailevatoron 138. Ailevatoron moment 142 is shown in a nose-up directionfor symmetric deflection(s) of ailevatoron(s) 138 that acts tosupplement reduced moment 136 up to a magnitude of rotation moment 130.

Conventionally, each of the moments shown in FIG. 1, may be analyzed asa moment about the center of gravity 110 due to the same forces shownand discussed for FIG. 1, but with a different moment arm. Thus, anymoment represented in FIG. 1 may have a comparable moment that iscalculated about center of gravity 110 instead of the rotation point ofaxis 106. Accordingly, although responsive to the same forces, momentscalculated about center of gravity 110 would have a different magnitude,due to the reference point shift and moment arm difference, from momentscalculated about axis 106 caused by the same forces.

Ailevatoron 138, may have a same shape, material, strength, range ofmotion, size, and be located in a same place on an aircraft as anaileron located on a particular aircraft model before replacinghorizontal stabilizer 116 with reduced horizontal stabilizer 132. In anembodiment, ailevatoron 138 may replace an aileron of the particularmodel with a flight control surface that has a different shape,material, strength, and/or size than the aileron. In an embodiment, alocation of ailevatoron 138 may be moved to a different location on wing126 than location of the aileron of the particular model beforereplacing horizontal stabilizer 116 with reduced horizontal stabilizer132. Ailevatoron 138 may be moved further aft, towards aircraft tail,from axis 106, via moving ailevatoron 138 further outboard on a sweptwing 126, than the location for the aileron originally on the particularmodel before replacing horizontal stabilizer 116 with reduced horizontalstabilizer 132.

Deflection of a trailing edge of ailevatoron 138 may be activated bysame actuator as aileron actuator originally on the particular modelbefore replacing horizontal stabilizer 116 with reduced horizontalstabilizer 132. In an embodiment, actuator for ailevatoron 138 may bechanged from aileron actuator to an ailevatoron actuator. Theailevatoron actuator may provide the technical effect of making a rateof movement of ailevatoron 138 faster and/or more precisely controlledthan a rate of movement previously capable for the aileron originally onthe particular model before replacing horizontal stabilizer 116 withreduced horizontal stabilizer 132.

Ailevatoron 138 may be commanded to displace from a position flush witha trailing edge of wing 126 and thereby generate force 140 thatgenerates ailevatoron moment 142. Force 140 may be an aerodynamic forceresultant from a drag and a lift caused by wing 126. Ailevatoron moment142 can be generated with a magnitude sufficient to supplement reducedmoment 136 up to equal rotation moment 130 required for takeoff.

Each wing 126 will include at least one ailevatoron 138 located along atrailing edge of each wing 126. Wing 126 trailing edge is the edge ofwing 126 closest to tail 118.

Generally, ailevatoron moment 142 is generated by a rotation of atrailing edge of ailevatoron 138 upward (TEU) and away from surface 102.The trailing edge is the edge closest to tail 118 of aircraft 100.Usually, the trailing edge of ailevatoron 138 on each left wing 126 willdeflect upward symmetrically and in equal magnitude with deflection ofthe trailing edge of ailevatoron 138 on right wing 126.

However, in an embodiment, if roll commands are being input at the timeailevatoron 138 trailing edge upward deflection is scheduled, it ispossible that deflection of the trailing edge of ailevatoron 138 on leftwing 126 may deflect upward asymmetrically relative to deflection of thetrailing edge of ailevatoron 138 on right wing 126 if such asymmetry isneeded to supplement roll control of aircraft 100. Hereinafter, anyreference to a deflection of ailevatoron 138 will Indicate deflection ofthe trailing edge of ailevatoron 138.

In an embodiment, an ailevatoron mixer, as shown without limitation, asailevatoron mixer 220 in FIG. 2, may act in series between and aileronmixer and an aileron/ailevatoron actuator with an activation, triggeredby a weight-on-wheels condition, that adds commands, for degree ofsymmetric ailevatoron 138 deflections that create ailevatoron moment142, onto existing roll commands to ailevatoron 138 from an aileronmixer. In an alternate embodiment, ailevatoron mixer outputs may bereceived into an aileron mixer and added in the aileron mixer and sentto the aileron/ailevatoron actuator.

The summing effect of roll commands and pitch commands through theailevatoron mixer 220 may result in an asymmetric defection of anailevatoron 138 on wing 126 on a left side of aircraft 100 relative towing 126 on a right side of aircraft 100. Thus, as a non-limitingexample, and as will be described in further detail later, if asrotation velocity occurs, a roll command to an aileron mixer exists thathas ailevatoron 138 on wing 126 on left side of aircraft 100 deflecteddownward 10 degrees, and ailevatoron 138 on wing 126 on right side ofaircraft 100 at a 10 degree upward deflection, then when a pitch commandto rotate the aircraft is received at rotation velocity, the ailevatoronmixer 220 will add, as a non-limiting example, 10 degrees upwarddeflection to ailevatoron 138 on each wing. Thus, at rotation andlift-off for aircraft 100, ailevatoron 138 on left wing of aircraft 100would be 0 degrees and ailevatoron 138 on right wing of aircraft 100would be deflected up 20 degrees.

In an alternate embodiment, an aileron mixer may be replaced byailevatoron mixer 220. In such an embodiment, ailevatoron mixer 220receives all roll commands as well as pitch rotation commands, and sumsthe commands for symmetric deployment of ailevatoron 138 to generateailevatoron moment 142 with roll commands to determine ultimate commandsfor deflection sent to actuators for ailevatoron 138 on each wing. Thus,as a non-limiting example, and as will be described in further detaillater, if as rotation velocity occurs, a roll command to ailevatoronmixer 220 exists that has ailevatoron 138 on wing 126 on left side ofaircraft 100 deflected trailing edge downward 1 degree, and ailevatoron138 on wing 126 on right side of aircraft 100 deflected trailing edgeupward 1 degree, then when a pitch command to rotate the aircraft isreceived in ailevatoron mixer 220 at rotation velocity, the ailevatoronmixer 220 will add, as a non-limiting example, 10 degrees upwarddeflection to ailevatoron 138 on each wing. Thus, at rotation andlift-off for aircraft 100, ailevatoron 138 on left wing of aircraft 100would be deflected up 9 degrees and ailevatoron 138 on right wing ofaircraft 100 would be deflected trailing edge upward 11 degrees.

Aircraft nose-up rotation at rotation speed is commanded with a nose-uppitch command that initially directs deflection of an elevator onreduced horizontal stabilizer 132. To maintain consistency in commandresponse between operation of a particular aircraft model withhorizontal stabilizer 116 and reduced horizontal stabilizer 132,generation of ailevatoron moment 142 may be programmed such that nose-uppitch response of aircraft 100 remains the same for similar commandinputs regardless of whether horizontal stabilizer 116 or reducedhorizontal stabilizer 132 were mounted on aircraft 100. Thus, to anoperator, the technical effect of ailevatoron 138 will not changerequired command inputs or pitch response of aircraft 100 to commandinputs.

In other words, the response and feel of aircraft 100 at rotation wouldbe seamless, or without difference to an operator of aircraft 100.Hence, generation of ailevatoron moment 142 that supplements reducedmoment 136 keeps rotation moment 130 generated equivalent betweenaircraft mounted with horizontal stabilizer 116 and reduced horizontalstabilizer 132, and thereby provides the enhanced technical effect ofincreased aircraft performance and fuel efficiency without requiring anyadditional inputs or training for an operator of aircraft 100.

Thus, a novel machine referred to as an ailevatoron mixer is designed toproduce the technical effect of commanding the ailevatoron 138 toproduce the force 140 required to supplement reduced moment 136 producedby reduced horizontal stabilizer 132 such that rotation moment 130produced by aircraft 100 with reduced horizontal stabilizer 132 may beequivalent in rate of application and magnitude to rotation moment 130produced by aircraft 100 with horizontal stabilizer 116.

With reference to FIG. 2, an illustration of aschematic representationof relative magnitudes of rotation moment control for an aircraft, andcomponents contributing thereto, is depicted in accordance with anillustrative embodiment. In this illustrative example, vertical axis 200represents magnitude (increasing upward along vertical axis 200) ofnose-up pitch moments acting on aircraft 100. Conventionally, magnitudesfor the moments may be calculated as acting about center of gravity 110of FIG. 1. Horizontal axis 202 represents a magnitude (increasing to theright along horizontal axis 202) of an aircraft airspeed, with rotationspeed (also referred to as V_(r)) marked by V_(R) line 204.

Level 206 represents a magnitude of rotation moment 130 about axis 106,in an aircraft 100 nose-up direction, required for aircraft 100 rotationat takeoff and initial climb of aircraft 100 away from surface 102.Curve 207 represents a magnitude of stab moment 124 produced byhorizontal stabilizer 116 up to rotation speed, V_(r) represented atV_(R) line 204. Curve 208 represents a magnitude of reduced moment 136produced by reduced horizontal stabilizer 132. Curve 210 representsailevatoron moment 142 produced by ailevatoron(s) 138 deflection(s),just prior to rotation speed 204, symmetrically upward away from surface102. Depending upon specific characteristics of aircraft 100, curve 207,curve 208, and curve 210 may be: curved, linear, linear with curvedportions, or curved with linear portions, so long as at V_(R) line 204,a magnitude of ailevatoron moment 142 represented by curve 210 plus amagnitude of reduced moment 136 represented by curve 208 equal therequired magnitude of rotation moment 130 represented by level 206.Generally, a magnitude for ailevatoron moment 142 at rotation (V_(R)line 204), will be in the range of 12-17% of the magnitude for reducedmoment 136 if the moments are computed about axis 106. Concurrently, amagnitude for ailevatoron moment 142 at V_(R) line 204, will be in therange of 14-20% of the magnitude for reduced moment 136 if the momentsare computed about center of gravity 110.

The distance from the point where curve 210 intersects horizontal axis202 to V_(R) line 204 represents an amount of time, deflection period209 of the ailevatoron(s) 138, which is dependent on acceleration ofaircraft 100. Similarly, the increase in velocity from V_(R) to thevelocity at the point where curve 214 intersects horizontal axis 202represents an amount of time for a washout of the symmetric deflectionof the ailevatoron(s) 138, washout period 211 for the ailevatoron(s)138.

In FIG. 2, deflection period 209 and washout period 211 may be shownexpanded for ease of viewing, and may not be shown in scale to eachother, or in scale to curve 207, curve 208, or curve 210. Likewise,relative magnitudes of curve 210 and 208 are not necessarily scaled tolevel 206 beyond representing the cumulative nature of the momentsindicated by the two curves.

Deflection period 209 and washout period 211 may each be dependent uponspecific characteristics of aircraft 100 and of environment for a giventakeoff, such as without limitation, thrust, altitude, air temperature,configuration of aircraft 100, location of center of gravity 110 onaircraft 100, as well as other aerodynamic and environmental factors. Inan embodiment, deflection period 209 and washout period 211 may be adifferent length from each other and may each be less than one second.In an embodiment, deflection period 209 and washout period 211 may be adifferent length from each other and may each be between 2 and 0.1seconds.

As discussed above, as soon as main gear 104 lifts off surface 102, andaircraft 100 is flying, aircraft pitch no longer rotates about axis 106,but about center of gravity 110. In flight, to control a pitch attitudefor aircraft 100, aerodynamic forces generated by wing 126 createmoments that act about center of gravity 110, and become the principalforces that must be counteracted by aerodynamic force 134 from reducedhorizontal stabilizer 132.

After V_(r), to the right of V_(R) line 204, curve 207 represents amagnitude, for a pitch moment acting about aircraft center of gravity110 after aircraft 100 lifts off from surface 102, required to set anattitude of aircraft 100 at a desired pitch for initial climb away fromsurface 102. After V_(r), to the right of V_(R) line 204, curve 210represents a magnitude of a nose-up moment, acting about center ofgravity 110, that is generated by the ailevatoron 138 as the upwarddeflection of ailevatoron 138 is removed, or “washed out”. After V_(r),to the right of V_(R) line 204, curve 208 represents a magnitude of amoment generated about center of gravity 110 by reduced horizontalstabilizer 132.

When airspeed increases to a magnitude represented by line 238, thencurves 208 and 207 join each other. When airspeed of aircraft 100increases above a magnitude represented by line 238, then reducedhorizontal stabilizer 132 becomes able to produce force 134 at amagnitude great enough to provide reduced moment 136 at magnitudesrequired for aircraft 100 to continue climb up away from surface 102without a supplement from ailevatoron moment 142 from ailevatoron 138.

Magnitudes, represented by curve 210, of moments produced by deflectionof ailevatoron 138 upward (away from surface 102) are a function ofairspeed 218 of aircraft 100 and an amount of degrees of deflection ofailevatoron 138. Deflection of ailevatoron 138 is commanded by novelailevatoron mixer 220. As illustrated by the schematic on the top halfof FIG. 2, ailevatoron mixer 220 commands a deflection for ailevatoron138 based upon inputs to ailevatoron mixer 220 of at least airspeed 218of aircraft 100, aircraft 100 status relative to surface 102,acceleration 242 of aircraft 100, yaw effects 244 and roll command 246,moments produced by reduced horizontal stabilizer 132 due to trim 224,and deflection 234 of elevator 226 on, reduced horizontal stabilizer132, as well as feedback on pitch rate 228 and pitch angle 230, andmoments produced by other aircraft components.

Aircraft 100 status relative to surface 102 may be provided toailevatoron mixer 220 by weight-on-wheels signal 222. Ailevatoron mixer220 contains filter 236, and will not pass pitch command 232 toailevatoron 138 unless each input into filter 236 is at a desired levelfor the component input. As an example, ailevatoron mixer 220 will notcommand upward deflection of ailevatoron 138 when weight-on-wheelssignal 222 tells ailevatoron mixer 220 that aircraft 100 main gear 104are not on surface 102.

when ailevatoron mixer 220 receives a change in weight-on-wheels signal222, from indicating aircraft 100 is on surface 102 to indicating thataircraft 100 is no longer on surface 102, ailevatoron mixer 222 commandsailevatoron 138 to “washout” 248 ailevatoron 138 symmetric deflection252 upward, and respond to roll command 246 for aircraft 100 in themanner of an aileron, as mounted and controlled on aircraft 100 beforereplacing horizontal stabilizer 116 with reduced horizontal stabilizer132. As shown by curve 214, washout 248 of symmetric upward deflectionis an elimination, within a designated period of time, of any addedsymmetric deflection 252 upward applied to ailevatoron(s) 138.

Reduced moment 136 produced by reduced horizontal stabilizer 132 is afunction of airspeed 218, acceleration 242, trim 224 of reducedhorizontal stabilizer 132, and a deflection of elevator 226 on reducedhorizontal stabilizer 132, as well as deflection 234 of elevator 226 ofaircraft 100. Scheduler 250 may be located within confines ofailevatoron mixer 220 as shown, or within a flight control computer incommunication with ailevatoron mixer 220.

Ailevatoron mixer 220 may be a part of and located within a flightcontrol computer. Deflection of elevator 226 is determined by pitchcommand 232 sent to elevator 226. Pitch command 232 sent to elevator 226is based upon desired rate 228 of a change in pitch angle 230 foraircraft 100 and a desired pitch angle 230 for aircraft 100 in additionto weight-on-wheels signal 222. Pitch command 232 is not processedthrough scheduler 250 to determine symmetric deflection 252 to be addedto aileron controller 240 input to mixer 256 to form command 254 todeflect ailevatoron 138 unless each condition sent into filter 236 iswithin a designated acceptable value for that condition. Acceptablevalues designated for each condition may be determined via valuesreceived by filter 236 of other conditions.

FIG. 2 and filter 236 are not; meant to depict a singular hardwareconfiguration for the embodiment, but rather illustrate a logic andoutcome of a configuration for filter 236 and ailevatoron mixer 220. Inall cases, ailevatoron mixer 220 only passes pitch command 232, alongwith values for each condition input into filter 236, on to scheduler250 when, each condition input is within an designated range. Asdepicted by FIG. 2, conditions input include at least trim 224,deflection 234 of elevator 226, airspeed 218 of aircraft 100,acceleration 242 of aircraft 100, pitch angle 230, pitch rate 228, andweight-on-wheels signal 222. As depicted by FIG. 2, when a givencondition is within a designated range (acceptable), filter 236 releasesa switch for the particular condition, from unacceptable position u,that prevents values of the particular condition from reaching scheduler250, down to position “a” below position “u”, so that information ofacceptable condition may be passed on to scheduler 250 when allconditions are within their respective designated range. As shown inFIG. 2, every switch for every condition is at unacceptable position“u”.

As a non-limiting example, if a command was given to immediatelyincrease pitch of aircraft 100 to rotate and takeoff, and weight was onwheels, and aircraft 100 was accelerating properly, but airspeed was 20knots below rotation speed, filter 236 would not allow the switch forairspeed to move from its “u” position to its “a” position, and thusscheduler 250 would net receive signals for each condition, and wouldnot send symmetric deflection 252 value into mixer 256. As anothernon-limiting example, if all conditions were within designated ranges,then all switches would be in acceptable “a” positions and all signalswould pass into scheduler 250 that determines magnitude of symmetricdeflection 252 to send to mixer 256. Magnitude of symmetric deflection252 determined by scheduler 250 based upon conditions of elevatordeflection 234 is shown in FIG. 3. As a further non-limiting example,when weight-on-wheels signal 222 changes from a condition of aircraft100 being on surface 102 to aircraft 100 not being on surface 102,washout 248 filter sends values into mixer 256 to washout any existingsymmetric deflection 252 over a designated time to a value of zero.Without limitation, designated time may be in a range from 2.0-0.2seconds.

Hence, activation of ailevatoron(s) 138 by ailevatoron mixer 220 isdynamic, and driven by dynamic symmetric deflection 252 determinationsfrom scheduler 250 based upon conditions input to scheduler 250. Thus,as a non-limiting example, degrees of symmetric deflection may vary asairspeed 218 of aircraft 100 changes. Filter 236 also functionsdynamically in real-time, such that acceptable ranges for any individualcondition input may change dependent upon current value of otherconditions input to filter 236. Filter 236 may be configured as programcode stored in a processor.

As discussed further below, filter 236 and ailevatoron mixer 220 may beformed via hardware, a specially programed code in a processor, ACSIcircuits, other equipment and methods described for FIG. 4, and/orcombinations thereof. In an embodiment, filter 236 may be within aflight control computer in communication with ailevatoron mixer 220. Inan embodiment, filter 236 and ailevatoron mixer 220 may each bepartitions within a flight control computer, when pitch command 232 isnot passed through scheduler 250 to create symmetry deflection 252, thenmixer 256 creates command 254 directly from input from aileroncontroller 240 without any additions.

Based upon a known configuration for components of a particular model ofaircraft, such as without limitation aircraft 100, for any givencombination of weight, location of center of gravity, and airspeed, foraircraft 100, moments produced by aircraft 100 components may be sensedby various sensors, or computed based on other sensed components andstates of aircraft 100. As a result, at rotation speed for aircraft 100,a magnitude for a moment required to rotate aircraft 100 to a pitchangle 232 required for takeoff (called takeoff attitude) rotation moment130, will be known as described above.

Hence, for a set rotation speed, and trim 224 of a set value for reducedhorizontal stabilizer 132, a particular amount of deflection 234 ofelevator 226 needed for rotation of aircraft 100 to the takeoff attitudemay be computed by a flight control computer for aircraft 100.Ailevatoron mixer 220 will be configured such that as elevator 226approaches deflection 234, ailevatoron mixer 220 will issue commands todeflect ailevatoron 138 upward on a schedule such that at deflection 234of elevator 226 ailevatoron 138 is fully deflected upward.

Ailevatoron mixer 220 is configured such that a magnitude and rate ofdeflection upward of ailevatoron 138 occurs such that a pitch angle 232for pitch rate 230 for a change in pitch angle 232 for aircraft 100occurs in a similar manner as would result for aircraft 100 whenconfigured with horizontal stabilizer 116. Conversely, once the aircraft100 lifts off from surface 102, filter 236 in ailevatoron mixer 220commands removal of upward deflection of ailevatoron 138 that was inputto the effect rotation of nosewheel 108 off of surface 102. Washout 248may be a filter that removes ailevatoron 138 trailing edge upwarddeflection, as shown at least by line 214 in FIG. 2. Washout 236 filteris configured to return ailevatoron 138 response and operation to thatof a traditional aileron on wing 126.

In some embodiments, ailevatoron 138 response to roll command 246 may beprogrammed by ailevatoron mixer 220 to differ from a response programmedfor traditional ailerons on aircraft 100 before reduced horizontalstabilizer 132, based upon changes in roll dynamic performance ofaircraft 100 due to reduced size and/or weight of reduced horizontalstabilizer 132 as compared to horizontal stabilizer 116. Adjustments toprevious aileron response after line 238 speed due to reduced horizontalstabilizer 132 may be referred to as adjusted aileron response.

Ailevatoron 138 upward deflection washes out from a maximum value forupward symmetric deflection 252 at rotation of aircraft 100, to no addedupward symmetric deflection 252, beyond aileron controller 240 responseto roll command 246 input as expected for an aircraft configured withoutailevatoron mixer 220. As illustrated, upward symmetric deflection 252decreases to decrease magnitudes of curve 210 (generated by ailevatoronmoment 142) as airspeed increases from rotation speed (V_(r) line 204)to an airspeed marked by vertical line 238. Vertical line 238 marks thespeed at which reduced horizontal stabilizer 132 has the authority toproduce pitch moments required to control the pitch of the aircraft, asshown by curve 208 meeting curve 207 in FIG. 2.

Further, ailevatoron mixer 220 can be configured such that aileronsmounted on wing 126 need no modification from a size and range of motionthat they are capable of for aircraft 100 with horizontal stabilizer116. Thus, in an embodiment, an aileron can become an ailevatoron byvirtue of being under control of ailevatoron mixer 220.

Where the original size and range of motion for the ailerons and onaircraft 100 is unchanged when horizontal stabilizer 116 is changed toreduce stabilizer 132 the installation of ailevatoron mixer 220 is whatchanges an aileron to become ailevatoron 138. Although FIG. 2, onlyshows a single ailevatoron 138 on an outboard (near wing tip—away fromfuselage) portion of wing 126, for wings that contain more than oneaileron, each aileron can be converted to an ailevatoron 138 by theaileron receiving commands from ailevatoron mixer 220, so long as theaileron is located aft of axis 106.

Thus the technical effect of ailevatoron mixer 220 and reducedhorizontal stabilizer 132 will be that, at rotation speed, aircraft 100will rotate to the same takeoff attitude at the same pitch rate asaircraft 100 would have with horizontal stabilizer 116 and noailevatoron mixer 220, such that pitch commands issued to ailevatoronmixer 220 may be unchanged from pitch commands that would be issued toan elevator controller on horizontal stabilizer 116 for aircraft 100. Inother words, an operator sending a command to rotate aircraft 100 to atakeoff attitude needs no additional training to issue commands tooperate aircraft 100 with ailevatoron mixer 220 and reduced stabilizer132, at least because the commands an operator would issue are the sameas would have been issued to aircraft 100 with horizontal stabilizer 116and no ailevatoron mixer 220.

Ailevatoron mixer 220 controls ailevatoron 138 by commands to anactuator of ailevatoron 138. In the case where ailevatoron 138 remainsthe same hardware as the original aileron (or ailerons as applicable) ofwing 126, ailevatoron mixer 220 may be configured to act as a supplementor addition to commands from a separate aileron controller 240 thatsends commands to the actuator of former aileron, now ailevatoron 138.Alternatively, ailevatoron mixer 220 may be configured to receivecommands from aileron controller 240 and process them with other inputsbefore sending commands to the actuator of ailevatoron 138, such thatailevatoron mixer 220 is the only unit sending commands to the actuatorof ailevatoron 138.

Aileron controller 240 and ailevatoron mixer 220 may be configured for“fly-by wire” (FBW) control systems, or for traditional cabled flightcontrol systems that are mechanically connected to a command inputdevice. A fly-by-wire (FBW) system for an aircraft is a system thatreplaces the traditional flight controls of an aircraft, which aremechanically connected to an input command device, with an electronicinterface.

For FBW, the input command device is not connected to the flight controlsurfaces, engines, or other systems by cables, linkages, or othermechanical systems, as in conventional aircraft. Instead, the movementsof flight controls are converted to electronic signals transmitted bywires, optical fibers, over an air-interface, or some combinationthereof.

The different components in a fly-by-wire system may communicate witheach other using different types of communications architectures. Forexample, some fly-by-wire systems use wires that connect the componentsdirectly to each other. In this example, multiple wires can be used toprovide redundant connections between the components.

In other examples, a fly-by-wire system may use a data bus, such asthose used in computer systems. The data bus may reduce the amount ofwiring between components. Wireless transmission of command signals mayalso be used.

For example, flight control computers in a fly-by-wire system usesignals to identify how to move the actuators for each flight controlsurface to provide the desired aircraft response to the movement of theflight controls. Further, flight control computers also may performfunctions without input from a pilot displacing a control yoke. In placeof commands generated by a mechanical displacement of a yoke, commandinputs may be generated by a pilot applying pressure onto a controlinput device, such as without limitation a control wheel steeringtransducer, a side stick, or a joystick. Further, commands may begenerated from other sources, such as without limitation, a flightcontrol computer, and/or a controller linked to the aircraft fromoutside the aircraft.

An aircraft with a fly-by-wire system can be lighter in weight than whenusing conventional controls. Electronic systems in a fly-by-wire systemrequire less maintenance as compared to flight control systems usingpurely mechanical systems and hydraulic systems.

Redundancy is present in fly-by-wire systems for aircraft. Multipleflight control modules in the fly-by-wire system may be used to generatecommands in response to receiving signals from the movement of flightcontrol-external sensing devices. The different components in afly-by-wire system may communicate with each other using different typesof communications architectures. For example, some fly-by-wire systemsuse wires that connect the components directly to each other. In thisexample, multiple wires can be used to provide redundant connectionsbetween the components.

In other examples, a fly-by-wire system may use a data bus, such asthose used in computer systems. The data bus may reduce the amount ofwiring between components.

Accordingly, in an embodiment with a FBW flight control system, aileroncontroller 240 may be fully replaced by ailevatoron mixer 220 thatreceives all signals that would have been intended for aileroncontroller 240. In such embodiments, ailevatoron mixer 220 wouldessentially subsume aileron controller function within ailevatoron mixer220 at a point inside ailevatoron mixer 220 just to the left of mixer256. Roll command 246 would input directly into ailevatoron mixer 220prior to mixer 256 and be processed by ailevatoron mixer 220.

Ailevatoron mixer 220 therefore, is configured to act throughout most offlight as would aileron controller 240, but ailevatoron mixer 220 isalso configured to deflect ailevatoron 138 upward to provide ailevatoronmoment 142 as needed, at relative magnitudes represented by lines 210,to produce required rotation moment 130. As mentioned above, when theprogram for aileron controller 240 with horizontal stabilizer 116 mayneed modification due to aerodynamic effects of smaller size of reducedhorizontal stabilizer 132, ailevatoron mixer 220 may send signals thatadjust commands of aileron controller to adjust for aerodynamic effectsof smaller size of reduced horizontal stabilizer 132. The aerodynamiceffects may include roll, pitch, or yaw differences that may vary invarious flight regimes for aircraft 100. Flight regimes may vary basedupon at least an altitude, airspeed, and a configuration of aircraft100. Configuration may include, without limitation, an extension statusof landing gear and/or flaps and/or flaperons, and/or leading-edgedevices.

At rotation, ailevatoron 138 upward deflection is biased to deflectailevatoron 138 on each wing 126 (left and right) on aircraftsymmetrically upward. However, ailevatoron mixer 220 also receives rollcommand 246 inputs and yaw effects 244 for aircraft 100, and if needed,can modulate deflection of ailevatoron 138 on one wing 126 to adeflection that may be asymmetric with deflection of ailevatoron 138 onwing 126 on opposite side of aircraft 100.

Alternatively, if it is desired to change a pitch response to particularrotation commands for aircraft 100, ailevatoron mixer 220 may beconfigured to receive the same commands that would have been issued forhorizontal stabilizer 116 and direct a different pitch response forreduced horizontal stabilizer 132. In other words, a pitch response ofaircraft 100 to takeoff commands given to aircraft 100 can be changedvia changes made to configuration and/or programming of ailevatoronmixer 220. Configuration and/or programming of ailevatoron mixer 220 mayinclude configuration of schedules of ailevatoron mixer 220.

As an additional feature, to enhance an ability of ailevatoron 138 toaffect a pitch rate and magnitude for aircraft 100, a size and/oravailable degrees of deflection upward for ailevatoron 138 can beincreased (relative to previous aileron design) during the aircraftdesign and manufacture, or as part of a retrofit, that becomes moreextensive modification of aircraft 100 than a replacement of onlyhorizontal stabilizer 116 with reduced horizontal stabilizer 132 inailevatoron mixer 220.

Changes to ailevatoron 138 may include a larger surface, an alternatelocation, and/or a greater upward deflection capability than provided byaileron design for aircraft 100 before reduced horizontal stabilizer132. Such increases can provide a greater magnitude for ailevatoronmoment 142 than could be produced by previous aileron design(size/shape/location/quantity) for aircraft 100 responding toailevatoron mixer 220 commands.

With reference to FIG. 3, FIG. 3 presents a graph that represents therelative magnitude of an upward deflection of a trailing edge of anailevatoron relative to a magnitude of a trailing edge of an elevatordeflection in a direction that generates a nose-up pitch for anaircraft, in accordance with an illustrative embodiment. In FIG. 3,vertical axis 302 of graph 300 represents deflection, as a percentage ofmaximum deflection capable, of elevator on reduced horizontal stabilizer132 in a direction that generates a nose-up pitch for aircraft 100.Horizontal axis 304 represents upward (away from surface 102(b))deflection, as a percentage of maximum deflection capable forailevatoron 138. As mentioned above the schedule would be commandedsymmetrically for ailevatoron 138 on each wing 126 of aircraft 100. Asdiscussed for FIG. 2, symmetric deflection 252 of ailevatoron 138 may bescheduled based upon several conditions, FIG. 3 gives an example ofvalues scheduled for symmetric deflection 252 based upon condition ofdeflection 234 of elevator 226.

For viewing clarity, the bottom of vertical axis 302 represents 50% ofelevator deflection towards a nose-up direction limited elevatordeflection. The precise relationships of deflections will depend uponflight control characteristics that currently exist for aircraft 100with horizontal stabilizer 116, or upon a desired design pitch responserotation characteristic for aircraft 100.

Thus the relative representations shown by graph 300 are presented as anexample of relative relationships that ailevatoron mixer 220 may beconfigured to produce. When aircraft 100 is at rotation speed, elevator226 deflection 234 is at percentage indicated by level 306, which issome percentage less than 100% full deflection in direction thatproduces a nose-up moment for aircraft 100. As a non-limiting example,level 306 may represent 93% of full deflection, or for another aircraftdesign, level 306 may represent 95% of full deflection, while on someaircraft level 306 may represent 99% of full deflection, or nearly fulldeflection of elevator 226.

As elevator deflection approaches level 306, at some level 308, ofelevator deflection, ailevatoron mixer 220 receives value of an elevatordeflection and initiates upward deflection of ailevatoron 130. Theprecise value for level 308 is determined by the desired rotationcharacteristics for aircraft 100. As mentioned above, level 308 may beselected to match the response to commands for horizontal stabilizer 116or may be selected at a different value to establish a new rotationcharacteristic for aircraft 100. Thus, particularly when ailevatoronmixer 220 is part of an FBW flight control system, rotationcharacteristics may even be programmed to vary based upon a crew incommand of aircraft 100 or a mission for a given flight of aircraft 100.In other words, a technical effect of ailevatoron mixer 220 is that itprovides an efficient mechanism to alter, if desired and/or required,the rotation characteristics for each flight of aircraft 100. Rotationcharacteristics may include a relationship between a magnitude of apitch command input to a pitch moment generated, a range 112 allowablefor center of gravity 110 at takeoff rotation—and thus a loading ofaircraft 100, rotation speed, and/or combinations thereof.

As a non-limiting example level 308 may indicate 80%, 85%, or some othervalue, of maximum deflection for elevator 226. Further, in anembodiment, and particularly so where ailevatoron mixer 220 is a part ofa fly-by-wire flight control system, a value for the airspeed 218acceleration 242 of aircraft 100 may be received by ailevatoron mixer220, and a value for level 308 may be determined by ailevatoron mixer220 based upon the acceleration 242 of airspeed 218 of aircraft 100.Thus, when thrust output from aircraft engines, environmentalconditions, tire inflation, or other factors, may influence acceleration242 of airspeed 216 as aircraft 100 approaches rotation speed, symmetricupward deflection of each ailevatoron 138 may begin at a percentage ofelevator 226 deflection 234 that allows ailevatoron 138 to reach level312 deflection value at rotation speed.

Curve 310 on graph 300 shows how ailevatoron mixer 220 schedulescommands for upward deflection of ailevatoron 138 based upon deflectionof elevator 226, such that as deflection of elevator 226 reaches level306, deflection of ailevatoron 138 reaches level 312, which approaches100% of full upward deflection. Curve 310 may be a line, a curve, and ora curve with a linear portion. Scheduler 250 may also use other chartsfor other conditions. For example, graph 300 may be used when all otherconditions are each at one particular stable value, or for some range ofacceptable values for some of the conditions.

Similar to level 308, the precise value for level 312 is determined bythe desired rotation characteristics for aircraft 100. As mentionedabove, level 312 may be selected to match the response to commands forhorizontal stabilizer 116 or may be selected at different value toestablish new rotation characteristics for aircraft 100.

As a non-limiting example, level 312 of FIG. 3 may indicate 95%, or 99%,or some other value, of maximum trailing edge upward deflection forailevatoron 138. Level 312 and/or 306 may each be less than 100% toallow for contingencies of nonstandard conditions that may suddenlyoccur at rotation speed, such as without limitation wind shifts orgusts, or a sudden loss of thrust by an engine of aircraft 100.

Additionally ailevatoron mixer 220 may be programmed to scheduleailevatoron 138 deflection upwards at just below 100% to generateailevatoron moment 142 component of rotation moment 130 that allows forsome additional upward deflection of ailevatoron 138 if needed toprovide some degree of asymmetric ailevatoron 138 deployment if neededto augment roll command 246 and/or yaw effects of aircraft 100 from justprior liftoff of aircraft 100 from surface 102 until just after liftoff.

Thus, ailevatoron mixer 220 provides a novel technical effect necessaryfor a process for reducing a size of horizontal stabilizer 116 for aparticular aircraft model. Ailevatoron mixer 220 provides a noveltechnical effect of augmenting a nose-up moment, reduced moment 136, forthe aircraft, provided by a horizontal stabilizer of a reduced size,reduced horizontal stabilizer 132, for the particular aircraft model,such that rotation moment 130 required to bring nose 128 of aircraft 100to a takeoff attitude at rotation speed. Ailevatoron mixer 220 and theprocess enabled thereby, provides a further technical effect of reducinga weight of the aircraft 100 by reducing a weight of the horizontalstabilizer mounted on the particular aircraft model, relative to aweight of horizontal stabilizer 116, via using a horizontal stabilizerof a reduced size, reduced horizontal stabilizer 132, for the particularaircraft model.

When aircraft 100 includes a fly-by-wire flight control system,ailevatoron mixer 220 may be a component of a fly-by-wire flight controlsystem. As such, ailevatoron mixer 220 may communicate with or be a partof a flight control computer. As such, ailevatoron mixer 220 may beconsidered specialized program code operating in a data processingsystem.

As part of a fly-by-wire flight control system, ailevatoron mixer 220may communicate and operate in conjunction with a wing body loadalleviation control logic such as described in U.S. Pat. No. 8,024,079,assigned to The Boeing Company, which is incorporated herein in itsentirety. In such a configuration, ailevatoron mixer 220 would replaceor augment aileron command mixer as depicted in FIG. 7 of U.S. Pat. No.8,024,079.

Similarly, ailevatoron mixer 220 may communicate and operate inconjunction with a direct lift control system such as without limitationthat described in U.S. Pat. No. 8,712,606 and U.S. Pat. No. 9,415,860,assigned to The Boeing Company. Accordingly, the features presented inU.S. Pat. No. 8,712,606 and U.S. Pat. No. 9,415,860, assigned to TheBoeing Company, are incorporated herein in their entirety.

Similarly, ailevatoron mixer 220 may communicate and operate inconjunction with yaw generating systems such as without limitation thosedescribed in U.S. Pat. No. 7,367,530, assigned to The Boeing Company.Accordingly, the features presented in U.S. Pat. No. 7,367,530, assignedto The Boeing Company, are incorporated herein in their entirety.

Turning now to FIG. 4, a diagram of a data processing system forimplementing an ailevatoron mixer 220 is depicted in accordance with anadvantageous embodiment. In this illustrative example, data processingsystem 400 includes communications fabric 402, which providescommunications between processor unit 404, memory 406, persistentstorage 408, communications unit 410, input/output (I/O) unit 412, anddisplay 414.

Processor unit 404 serves to execute instructions for software that maybe loaded into memory 406. In an embodiment, processor unit 404 mayrepresent ailevatoron mixer 220 and may be a set of one or moreprocessors or may be a multi-processor core, depending on the particularimplementation. In an embodiment, processor unit 404 may represent aflight control computer and may be a set of one or more processors ormay be a multi-processor core, depending on the particularimplementation, of which ailevatoron mixer 220 may be a component. In anembodiment, processor unit 404 may represent a flight control system andmay be a set of one or more processors or may be a multi-processor core,depending on the particular implementation, of which ailevatoron mixer220 may be a component.

Further, processor unit 404 may be implemented using one or moreheterogeneous processor systems in which a main processor is presentwith secondary processors on a single chip. As another illustrativeexample, processor unit 404 may be a symmetric multi-processor systemcontaining multiple processors of the same type.

Memory 406, in these examples, may be, for example, a random accessmemory or any other suitable volatile or non-volatile storage device.Persistent storage 408 may take various forms depending on theparticular implementation. For example, persistent storage 408 maycontain one or more components or devices. For example, persistentstorage 408 may be a hard drive, a flash memory, a rewritable opticaldisk, a rewritable magnetic tape, or some combination of the above. Themedia used by persistent storage 408 also may be removable. For example,a removable hard drive may be used for persistent storage 408.

Communications unit 410, in these examples, provides for communicationswith other data processing systems or devices. In these examples,communications unit 410 is a work interface card. Communications unit410 may provide communications through the use of either or bothphysical and wireless communications links.

Input/output unit 412 allows for input and output of data with otherdevices that may be connected to data processing system 400. Forexample, input/output unit 412 may provide a connection for user inputthrough a keyboard and mouse. Further, input/output unit 412 may sendoutput to a printer. Display 414 provides a mechanism to displayinformation to a user.

Instructions for the operating system and applications or programs arelocated on persistent storage 408. These instructions may be loaded intomemory 406 for execution by processor unit 404. The processes of thedifferent embodiments may be performed by processor unit 404 usingcomputer implemented instructions, which computer may be located in amemory, such as memory 406. These instructions are referred to asprogram code, computer usable program code, or computer readable programcode that may be read and executed by a processor in processor unit 404.The program code in the different embodiments may be embodied ondifferent physical or tangible computer readable media, such as memory406 or persistent storage 408.

Program code 416 is located in a functional form on computer readablemedia 418 that is selectively removable and may be loaded onto ortransferred to data processing system 400 for execution by processorunit 404. Program code 416 and computer readable media 418 form computerprogram product 420 in these examples. In one example, computer readablemedia 418 may be in a tangible form, such as, for example, an optical ormagnetic disc that is inserted or placed into a drive or other devicethat is part of persistent storage 408 for transfer onto a storagedevice, such as a hard drive that is part of persistent storage 408. Ina tangible form, computer readable media 418 also may take the form of apersistent storage, such as a hard drive, a thumb drive, or a flashmemory that is connected to data processing system 400. The tangibleform of computer readable media 418 is also referred to as computerrecordable storage media. In some instances, computer readable media 418may not be removable.

Alternatively, program code 416 may be transferred to data processingsystem 400 from computer readable media 418 through a communicationslink to communications unit 410 and/or through a connection toinput/output unit 412. The communications link and/or the connection maybe physical or wireless in the illustrative examples. The computerreadable media also may take the form of non-tangible media, such ascommunications links or wireless transmissions containing the programcode.

The different components illustrated for data processing system 400 arenet meant to provide architectural limitations to the manner in whichdifferent embodiments may be implemented. The different illustrativeembodiments may be implemented in a data processing system includingcomponents in addition to or in place of those illustrated for dataprocessing system 400. Other components shown in FIG. 4 can be variedfrom the illustrative examples shown.

As one example, a storage device in data processing system 400 is anyhardware apparatus that may store data. Memory 406, persistent storage408 and computer readable media 418 are examples of storage devices in atangible form.

In another example, a bus system may be used to implement communicationsfabric 402 and may be comprised of one or more buses, such as a systembus or an input/output bus. Of course, the bus system may be implementedusing any suitable type of architecture that provides for a transfer ofdata between different components or devices attached to the bus system.Additionally, a communications unit may include one or more devices usedto transmit and receive data, such as a modem or a network adapter.Further, a memory may be, for example, memory 406 or a cache such asfound in an interface and memory controller hub that may be present incommunications fabric 402.

When ailevatoron mixer 220 is a part of or in communication with afly-by-wire flight control system, commands sent from ailevatoron mixer220 to ailevatoron 138 may be scheduled in response to commands input tonot only reduced horizontal stabilizer 132, but to other flight controlsas well, such as without limitation flaps, flaperons, rudders, and/orspoilers on aircraft 100. Additionally, commands sent from ailevatoronmixer 220 to ailevatoron 138 may be scheduled in response to currentpositions of each flight control surface of aircraft 100. Additionally,ailevatoron mixer 220 may send commands not only to ailevatoron 138, butto a controller and/or a mixer of any other flight control surface onthe aircraft. It is understood that commands sent to flight controlsurface may be sent to a respective actuator for the respective flightcontrol surface.

FIG. 5 is an illustration of a functional schematic diagram of anailevatoron mixer control and logic, in accordance with an advantageousembodiment. Functional schematic 500 illustrates ailevatoron mixer 220control integration with aircraft 100 that may contain a fly-by-wireflight control system. Ailevatoron mixer 220 communicates with flightcontrol system 502. Flight control system 502 may contain (as shown inFIG. 5), receive, be in communication with, and/or compute at leastrotation speed criteria 504 and flight control inputs 506 and maycontain non-linear filters 508 and linear filters 510 and/or a flightcontrol position 512, that each may form a part of status of aircraft100 analyzed by scheduler 250 in ailevatoron mixer 220 to determineoutput of ailevatoron mixer 220 to ailevatoron 138 and/or other flightcontrols.

Rotation speed criteria 504 may include, without limitation: anindicated airspeed for rotation, a trim setting for reduced horizontalstabilizer 132 for rotation, an engine thrust, a pressure altitude, aweight of aircraft, a weight-on-wheels signal 222, a configuration ofaircraft 100, a location of center of gravity 110 (cg) of aircraft 100,a length of surface 102, a slope of surface 102, a rotation rate foraircraft 100, and/or a rotation pitch target for aircraft 100. Pitchcontrol input 506 may be from a flight control device inside theaircraft, or from an autopilot command, and/or from a datalink input toaircraft 100 and/or ailevatoron mixer 220.

Linear filters 510 may include, without limitation, notch filters tominimize structural coupling. Linear filters 510 and non-linear filters508 filter out (or minimize) high gain feedback commands from wingstructural modes above one hertz. The feedback allows compensations toscheduled commands for lift and pitching moment changes due to controlinputs and/or atmospheric disturbances.

In one embodiment, a moving time window associated with the set oflinear filters 510 and non-linear filters 508 may be added to blockselected frequency signals. The time can be adjusted for a fasterresponse time or to block higher frequencies. The time may be a functionof aircraft state, flight conditions, structural modes frequencies,and/or weight of aircraft 100. Ailevatoron mixer 220 schedule inputs areschedules for extending and retracting control surfaces to supplementreduced moment 136 generated by reduced horizontal stabilizer 132.

Ailevatoron mixer 220 scheduler 250 may, without limitation, receiveinputs from or send output to other flight controls in addition toailevatoron 138. As a reminder, discussions in regard to ailevatoron 138are representative of each ailevatoron 138 on each wing, and if soinstalled, multiple ailevatorons on each wing.

Ailevatoron mixer 220 may communicate with other flight control mixersbased upon: inboard spoiler retraction and roll check back ailevatoronschedules 514, flap and/or flaperon direct lift and roll checkbackailevatoron schedules 516, ailevatoron 138 trailing edge up(TEU)/trailing edge down (TED) direct lift and roll checkbackailevatoron schedules 518, outboard spoiler extend/retract & rollcheckback ailevatoron schedules 520, and elevator pitch control lawreconfigure and crossfeed compensation 522. Ailevatoron mixer 220scheduler 250 selects a schedule for determining a magnitude forsymmetric deflection 252 based on a combination of aircraft 100dynamics, aircraft 100 current state, and aircraft 100 payload.

The ailevatoron mixer 220 utilizes scheduler 250 to allow for minimizingreduced moment 136 as compared to stab moment 124 for any requiredrotation moment 130. Ailevatoron 138 TEU/TED direct lift and rollcheckback ailevatoron schedules 518 may include one or more schedulesfor commanding ailevatoron(s) 138 symmetric trailing edge up deployment(TEU) deployment to generate force 140 and thus ailevatoron moment 142.Outboard spoiler extend/retract and roll checkback ailevatoron schedules520 may include one or more schedules for commanding outboard spoilersto extend and/or retract to maximize ailevatoron moment 142 whilemaintaining commanded roll control.

Elevator pitch control law reconfigure and crossfeed compensation 522may include schedules for feedback and crossfeed to one or moreelevators via an elevator command mixer 524 to compensate forailevatoron moment 142 and pitching moment changes. Ailevatoron mixer220 communication with other flight controls may be via: outboardspoiler command mixer 526 flap and/or flaperon command mixer 528, and/orinboard spoiler command mixer 530. Each communication and schedule maybe further adjusted by a respective gain schedule as shown in FIG. 5.

An ailevatoron mixer 220 may also be configured to the same effect foran aircraft without fly-by-wire controls. In an aircraft withconventional cable controls, connected to conventional pilot yoke inputdevices, ailevatoron mixer 220 may still receive trim setting of reducedhorizontal stabilizer 132, a weight-on-wheels signal 222 and aircraft100 airspeed inputs and recognize a rotation pitch command to theelevators, and upon recognizing elevator deflection reaching level 308,mechanically blending in a deflection upward, away from the surface 102,ailevatoron 138 along a schedule such as indicated by graph 300.

Similarly, in non-fly-by-wire aircraft, ailevatoron mixer 220 may havemechanical mixer valves to incorporate roll control inputs to aircraft100 into deflection of ailevatoron 138 to an effect similar to thatdescribed above. It is understood that the replacement of electronicprocessing for schedules of ailevatoron mixer 220 with mechanical mixervalves and cables will increase mechanical component requirements of anailevatoron mixer 220 system, and thus an overall weight, as compared toa fly-by-wire configuration.

Thus, described above is at least a novel machine configured to reduce asize of horizontal stabilizer 116 for a particular aircraft model whilegenerating and sustaining a nose-up moment, required for takeoff, forthe particular aircraft model loaded at a combination of a gross takeoffweight 122 and a center of gravity 110 located at a forward locationallowed in range 112 of allowable center of gravity locations, thatresults in a maximum required rotation moment 130, such that the machinecomprises ailevatoron mixer 220 configured to symmetrically deflectailevatoron(s) 138, located aft of an axis 106 of contact of main gear104 of aircraft 100, away from a surface 102. An embodiment of themachine also includes a flight control computer that comprises programcode fixed in a non-transitory medium configured to blend symmetricdeflection of a trailing edge, of ailevatoron(s) 138 away from surface102, based upon at least: deflection of elevator 226 on reducedhorizontal stabilizer 132, an airspeed, and a weight-on-wheels signal,of the particular aircraft model.

With reference now to FIG. 6, FIG. 6 is an illustration of a flowchartfor operations performed by one embodiment for reducing a size of ahorizontal stabilizer for a particular aircraft model, in accordancewith an advantageous embodiment.

Hence, process 600 may begin when an aircraft flight control systemreceives a signal that an aircraft is on a takeoff surface (operation602). The signal may come from a weight-on-wheels sensor, or anyappropriate sensor that recognizes that aircraft 100 main gear 104 aresupporting weight of aircraft 100 on surface 102. Flight control systemmay include a flight control computer, and/or a controller and/or amixer and/or an actuator unit for each and/or for any flight control inthe flight control system.

In process 600, the flight control system may receive a command torotate the aircraft to a takeoff pitch attitude on the takeoff surface(operation 604). The command may direct deflecting elevator 226 locatedon reduced horizontal stabilizer 132.

Upon an elevator deflection reaching a designated level, an ailevatoronmixer commands deflecting a trailing edge of each ailevatoronsymmetrically upward, away from the takeoff surface, to a designatedlevel (operation 606). Designated level 312 may be 100% deflection, orsome percentage less than 100%, based upon a schedule for commandingdeflection of ailevatoron 138 relative to deflection of elevator 226.

Responsive to the aircraft flight control system no longer receiving thesignal that the aircraft is on the takeoff surface, the ailevatoronmixer removes commands deflecting the trailing edge of each ailevatoronsymmetrically upward, away from the takeoff surface, to the designatedlevel (operation 608). Removing commands deflecting a trailing edge ofeach ailevatoron 138 symmetrically upward, away from surface 102, to adesignated level 312, may be scheduled as a blended reduction relativeto deflection of elevator 226, as illustrated by FIG. 2.

Thereafter, in process 600, and responsive to commands to roll theaircraft, the ailevatoron mixer commands each ailevatoronasymmetrically, and the ailevatoron mixer does not command deflectingany ailevatoron in response to commands to control the pitch of theaircraft (operation 610). As described above in detail, the ailevatoronmixer asymmetric commands to the ailevatoron will be comparable tocommands to an aileron on a particular aircraft model by an aileronmixer or aileron controller before the particular aircraft model had areduced horizontal stabilizer and the ailevatoron mixer installed. Thismay be achieved by the aircraft retaining aileron mixer commandsreaching the ailevatoron without addition deflection commanded by theailevatoron mixer, or by an aileron mixer and/or controller beingsubsumed within or replaced by the ailevatoron mixer. In other words,without limitation. FIG. 6 relates at least to a process and machine forreducing a drag component of a horizontal stabilizer on an aircraft.

Referring now to FIG. 7, FIG. 7 is an illustration of a flowchart foroperations performed by one embodiment for reducing a size of ahorizontal stabilizer for a particular model of an aircraft. Theoperations are represented by flowchart 700 that shows reducing a sizeof horizontal stabilizer 116 for a particular model of aircraft 100creating a reduced horizontal stabilizer (operation 702) and replacinghorizontal stabilizer 116 with reduced horizontal stabilizer 132.

Operations performed by embodiment shown by flowchart 700 continue byaugmenting a moment produced by the reduced horizontal stabilizer 132,known as reduced moment 136 (operation 704) via a deflection, away fromsurface 102, of ailevatoron 138. In an embodiment, ailevatoron 138 maybe an aileron already mounted on aircraft 100, under the control of theailevatoron mixer 220, such that each aileron that is located aft ofaxis 106 functions as ailevatoron 138, configured for deflectingsymmetrically, with ailevatoron 138 on wing 126 mounted on other side ofaircraft 100, away from surface 102 (operation 706).

Thus, a technical effect of ailevatoron mixer 220 is to utilize andcontrol the previously mounted ailerons, newly as ailevatoron (s) 138 onaircraft 100. Additionally, each previously mounted aileron may also beredesigned and/or replaced with newly sized and/or configuredailevatoron 138 tailored for aircraft 100 with reduced horizontalstabilizer 132.

A further technical effect of reducing a size of horizontal stabilizer116 on aircraft 100 will be to reduce a weight of aircraft 100 via areduced weight of smaller reduced horizontal stabilizer 132 relative tohorizontal stabilizer 116. Reduction in aircraft weight provides,without limitation, the technical effect of improving performance ofaircraft 100 with reduced horizontal stabilizer 132 as compared at leastto aircraft 100 with horizontal stabilizer 116. Without limitation,improved performance may include increasing: a range, an operatingceiling, a takeoff payload, a capability to accelerate, and/or a fuelefficiency; as well as reducing: a drag, a rotation speed, and/or atakeoff distance required, for aircraft 100.

Deflecting each ailevatoron 138 upward away from surface provides thetechnical effect of generating a force 140 located aft of axis 106,where main gear 104 of aircraft 100 contact the surface 102 (operation708). Ailevatoron mixer 220 receives a value for the degree ofdeflection 234 of a trailing edge of elevator 226 on reduced horizontalstabilizer 132, and based upon the value and a programmed schedule,commands symmetrically deflecting a trailing edge of each ailevatoron138 upward away from surface 102 (operation 710). The programmedschedule of the ailevatoron mixer 220 may be an algorithm in a computerprogram within a processor.

Ailevatoron mixer 220 may be considered a component within a flightcontrol computer and/or a flight control system. Alternately,ailevatoron mixer 220 program schedule may be mechanically programmedinto ailevatoron mixer 220 that mechanically blends commands fordeflection of a trailing edge of elevator 226 with ailevatoron 138deflection while aircraft main gear 104 contact the runway, surface 102.

Flowchart 700 may also include deflecting symmetrically, away fromsurface 102, a trailing edge of each ailevatoron 138 on aircraft 100based upon a degree of deflection of elevator 226 on reduced horizontalstabilizer 132 according to a determination in ailevatoron mixer 220 andat least one of: a weight-on-wheels signal 222, an airspeed 218 ofaircraft 100, and an acceleration rate of the aircraft (operation 712).

Further, operations performed by embodiment shown by flowchart 700 mayinclude basing a size, of reduced horizontal stabilizer 132 for theparticular aircraft model, for a rotation speed based upon an allowablecombination of gross takeoff weight and a forward center of gravity thatresults in a maximum nose-down moment (operation 714). An allowablecombination includes a gross takeoff weight, and a center of gravitylocation toward a forward end of a range, that are allowed by governmentregulation and/or certification, and/or aircraft manufacturer'soperating limits.

Although reduced horizontal stabilizer 132 may be designed into aircraft100 before manufacturing begins, reducing the size of the horizontalstabilizer for a particular aircraft model, such as without limitation,horizontal stabilizer 116 for aircraft 100 100, may occur by replacingthe horizontal stabilizer with the reduced horizontal stabilizer, suchas without limitation, horizontal stabilizer 116 with reduced horizontalstabilizer 132, after manufacturing the particular aircraft model withthe horizontal stabilizer, during at least one of: a modification, areconfiguration, a refurbishment, other maintenance, and otherservicing, of the particular aircraft model (operation 716).

Referring now to FIG. 8, FIG. 8 is an illustration of a flowchart foroperations performed by one embodiment for increasing the fuelefficiency for a particular model of an aircraft. Operations performedby embodiment shown by flowchart 800 begin via reducing a size of ahorizontal stabilizer for a particular aircraft model (802). Applicationof the ailevatoron mixer 220 and components and processes associatedtherewith for a particular aircraft model may result in reductions ofdrag during flight produced by reduced horizontal stabilizer 132 in therange of 5-10% for the aircraft model as compared to drag produced byhorizontal stabilizer 116 used before replacing horizontal stabilizer116 with reduced horizontal stabilizer 132. The reduced drag may beprofile drag. Still further size reduction for reduced horizontalstabilizer 132 and drag reductions may be possible when coupled with thenovel machine and process of U.S. patent application Ser. No.14/921,841, from the inventor of this application, which is fullyincorporated by reference herein. The reduced drag may be profile drag.Such further reductions may be due at least to a reduction in sizingrequirements that are currently driven by nose-down pitch controlrequirements, such as found without limitation, in 14 CFR § 25, foraircraft stall response handling.

Operations performed by embodiment shown by flowchart 800 continue byaugmenting a nose-up moment, for the particular aircraft model, providedby a reduced horizontal stabilizer, and reducing the weight of thehorizontal stabilizer for the particular aircraft model via the reducedhorizontal stabilizer (operation 804). Reducing a size required forhorizontal stabilizer 116 may have the technical effect not only ofreducing the weight of reduced horizontal stabilizer 132 and aircraft100 based upon use of similar materials, but may also reduce the weightbecause the reduced size may allow for construction of reducedhorizontal stabilizer 132 with a type of support and/or a type ofmaterials that weighs less than those used for horizontal stabilizer116. The reduced weight support and materials needed for reducedhorizontal stabilizer may result from changes in both aerodynamic andgravitational loading on the reduced horizontal stabilizer 132. Hence,the novel machine and process described herein may result in a reducedhorizontal stabilizer for a particular model of an aircraft that weighsless than a horizontal stabilizer required for the particular aircraftwithout utilizing the ailevatoron mixer 220 and related componentsand/or processes described herein. Without limitation, reducedhorizontal stabilizer 132 may weigh 10% less than horizontal stabilizer116.

Operations performed by embodiment shown by flowchart 800 continue byaugmenting a nose-up moment provided by the reduced horizontalstabilizer 132 via deflecting, away from surface 102, a trailing edge ofailevatoron 138 on aircraft 100 (operation 806). Operations continue byaugmenting the nose-up moment via generating a force, via the deflectionthe ailevatoron(s) 138 being located aft of axis 106 of contact of maingear 104 of aircraft 100 on surface 102 (operation 808).

Operations performed by the embodiment shown by flowchart 800 continueby deflecting symmetrically, away from surface 102, ailevatoron(s) 138on aircraft 100 based upon a degree of deflection, of elevator 226 onreduced horizontal stabilizer 132, according to a determination inailevatoron mixer 220(operation 810). Determination in may be based uponfactors described above. Ailevatoron mixer 220 may be in communicationwith or part of a flight control computer for a fly-by-wire flightcontrol system. In an embodiment, ailevatoron mixer 220 may be amechanical device that receives mechanical inputs from input device tocontrol pitch of aircraft 100 when on surface 102. As used herein,surface 102 may be a runway, or any surface used by aircraft 100 fortakeoff roll, rotation, and lift off.

Operations performed by an embodiment shown by flowchart 800 continue bydeflecting symmetrically, away from a runway, each ailevatoron on theparticular aircraft model based upon a degree of deflection of anelevator on the reduced horizontal stabilizer according to adetermination by an ailevatoron mixer considering at least: aweight-on-wheels signal, and an airspeed of the particular aircraftmodel (operation 812). Operations continue by basing the size of thereduced horizontal stabilizer, for a maximum nose-down moment for theparticular aircraft model, resulting from a particular combination ofgross takeoff weight and a maximum allowable forward center of gravity,upon a rotation speed for the particular aircraft model (operation 814).

Operations performed by an embodiment shown by flowchart 800 continue byreducing the size of the horizontal stabilizer 116 for a particularaircraft model by replacing horizontal stabilizer 116, aftermanufacturing the particular aircraft model using the horizontalstabilizer 116, with reduced horizontal stabilizer 132 during at leastone of: a modification, a reconfiguration, a refurbishment, othermaintenance, and other servicing, of the particular aircraft model(operation 816).

Additionally, an embodiment may at least provide operations forsupplementing reduced moment 136 generated by reduced horizontalstabilizer 132 for aircraft 100. Use of reduced horizontal stabilizer132 may provide the technical effect of reducing a takeoff airspeed, orincreasing a takeoff payload, for aircraft 100, at least via a reductionin an empty operating weight of aircraft 100.

Additionally, the process and machine described above provide thetechnical effect of supplementing a pitch moment produced by a givenhorizontal stabilizer for a given aircraft. Further, the ailevatoronmixer 220 may provide a means for providing a stall recovery means foraircraft 100 that may allow for further reduction of a size of thehorizontal stabilizer.

Additional technical effects also result from ailevatoron mixer 220adding symmetric deflection 252 into command 254 for activation ofailevatoron 138. The additional nose-up moment generated by variousdegrees of symmetric deflection 252 of ailevatoron 138 can also beapplied to provide mitigation of some failure modes at rotation speedfor aircraft 100, and/or expand a range for allowable takeoff trimpositions for a horizontal stabilizer of any given size.

As a non-limiting example, FIG. 2 can also be used to represent thebenefit of ailevatoron 138 symmetric deflection 252 when a horizontalstabilizer trim 224 is in a position beyond its normal operating rangefor an aircraft 100 without an ailevatoron mixer 220. If aircraft 100did not have a retrofit of reduced horizontal stabilizer 132, but didhave ailevatoron mixer 220 installed, then horizontal stabilizer 116could use a takeoff trim 224 setting outside of a previously allowablerange, such that during takeoff roll, horizontal stabilizer 116 at trim224 setting outside of a previously allowable range would producemagnitudes of a nose-up rotation moment similar to curve 208 in FIG. 2instead of the magnitudes shown for curve 207 that represents magnitudesfor horizontal stabilizer 116 with trim 224 at a previously normalsetting for takeoff. At rotation speed, ailevatoron 138 symmetricdeflection 252 would provide the additional nose-up moment magnitudeneeded to meet the required level 206. Thus, ailevatoron mixer 220 andailevatoron 138 not only provide the technical effect of allowing areduction in a size of horizontal stabilizer 116, but also allow for anexpansion of allowable takeoff trim 224 settings for horizontalstabilizer 116—or for reduced horizontal stabilizer 132. Even withoutchanging a size of horizontal stabilizer 116, changing a setting fortrim 224 of horizontal stabilizer 116 for takeoff can reduce dragproduced by horizontal stabilizer 116 during takeoff, and thus improveperformance and/or reliability for aircraft 100 via reducing either atakeoff distance required and/or a thrust required for takeoff.

Further, both technical effects can be achieved in varying degreessimultaneously. With ailevatoron mixer 220, a reduction in a size ofhorizontal stabilizer 116 to some intermediate size, between horizontalstabilizer 116 and reduced horizontal stabilizer 132, would allow forthe intermediate size horizontal stabilizer to operate with an expandedrange of allowable takeoff trim 224, as compared to operation of theintermediate size horizontal stabilizer without ailevatoron mixer 220.

Similarly, ailevatoron mixer 220 provides the technical effect ofaffording protections for certain failure modes of certain aircraftsystems during takeoff. As shown in FIG. 3, when level 312 designed forsymmetrical deflection 252 for a given rotation speed and trim 224setting lies below a value of 100%, then some extra amount of nose-upmoment may be generated by ailevatoron 138 above that shown by theintersection of curve 210 and V_(r) line 204 in FIG. 2. Hence, in theevent of some failure mode, such as without limitation, a mis-set trim224 position for reduced horizontal stabilizer 132 (or horizontalstabilizer 116 on an aircraft with ailevatoron mixer 220 but no reducedhorizontal stabilizer 132) and/or a jammed elevator 226 that inhibitsachieving desired deflection 234 required to produce planned takeoffrotation performance indicated by curve 208 (or curve 207) in FIG. 2,then, ailevatoron mixer 220 could direct an additional symmetricdeflection 252 (shown as available by the distance between level 312 and100% in FIG. 3) to increase the magnitude of 210 produced by ailevatoronmoment 142 being provided by ailevatoron 138. Thus, even if curve 203(or 207) intersected Vr line 204 at a lower level, ailevatoron mixer 220could provide the technical effect whereby curve 210 could be furtheraugmented to a higher level above original design to compensate for thedrop in magnitude of curve 208 (or 207).

Alternatively, a takeoff failure mode may alter FIG. 2 by moving themagnitude of level 206 upward. As a non-limiting example, a draggingbrake/deflated tire or a takeoff thrust below planned magnitude maychange rotation moment 130 and thus alter FIG. 2 by moving the magnitudeof level 206 upward. As above, ailevatoron mixer 220 could direct anadditional symmetric deflection 252 (shown as available by the distancebetween level 312 and 100% in FIG. 3) to increase the magnitude of 210produced by ailevatoron moment 142 being provided by ailevatoron 138.Thus, even if level 206 rises above a planned magnitude, ailevatoronmixer 220 could provide the technical effect whereby curve 210 could befurther augmented to a higher level above original design to compensatefor the rise of level 206 due to some takeoff failure condition onaircraft 100. Thereby, ailevatoron mixer 220 provides for enhancednose-up pitch authority for aircraft 100 at rotation that providesmitigation of failure modes within aircraft 100 components that affectrotation moment 130 for aircraft 100, as well as, improving aircraft 100performance via allowing a reduction in size of horizontal stabilizer116, and/or expanded allowable trim settings for horizontal stabilizer116 or reduced horizontal stabilizer 132.

Embodiments of the disclosure may be described in the context of theaircraft manufacturing and service operations 900 as shown in FIG. 9 foraircraft 100 as shown in FIG. 1. FIG. 9 shows a diagram illustratingoperations of an embodiment for an aircraft manufacturing and servicemethod, depicted in accordance with an advantageous embodiment. Duringpre-production, operations for aircraft manufacturing and serviceoperations 900 may include specification and design 902 of aircraft 100in FIG. 1 and material procurement 904.

During production, component and subassembly manufacturing 906 andsystem integration 908 of aircraft 100 in FIG. 1 takes place.Thereafter, aircraft 100 in FIG. 1 may go through certification anddelivery 910 in order to be placed in service 912. While in service by acustomer, aircraft 100 in FIG. 1 is scheduled for routine maintenanceand service 914, which may include modification, reconfiguration,refurbishment, and other maintenance or service.

Each of the processes of aircraft manufacturing and service operations900 may be performed or carried out by a system integrator, a thirdparty, and/or an operator. In these examples, the operator may be acustomer. For the purposes of this description, a system integrator mayinclude, without limitation, any number of aircraft manufacturers andmajor-system subcontractors; a third party may include, withoutlimitation, any number of venders, subcontractors, and suppliers; andwithout limitation an operator may be an airline, leasing company,military entity, service organization, and so on.

Machines and processes embodied herein may be employed during any one ormore of the stages of aircraft manufacturing and service method 900 inFIG. 9. For example, components or subassemblies produced in componentand subassembly manufacturing 906 in FIG. 9 may be fabricated ormanufactured in a manner similar to components or subassemblies producedwhile aircraft 100 is in service 912 in FIG. 9.

Also, one or more apparatus embodiments, method embodiments, or acombination thereof may be utilized during production stages, such ascomponent and subassembly manufacturing 906 and system integration 908in FIG. 9, for example, without limitation, by substantially expeditingthe assembly of or reducing the cost of aircraft 100. Similarly, one ormore of apparatus embodiments, method embodiments, or a combinationthereof may be utilized while aircraft 100 is in service 912 or duringmaintenance and service 914 in FIG. 9.

Additionally, the illustrative embodiments recognize and take intoaccount one or more different considerations. For example, theillustrative embodiments recognize and take into account that the use ofbuses, such as those used in computers, is becoming more common inaircraft. For example, special flight control programs in a computerprocessor may send commands to a special actuator control program in aprocessor that controls a device in the aircraft. An actuator controlprogram may control, for example, a flight control surface, an engine,or some other suitable device in the aircraft that may affect a changein pitch attitude or rate of an aircraft.

The illustrative embodiments also recognize and take into account that abus may be a parallel bus or a serial bus. When a parallel bus is used,units of data, such as a word, may be carried on multiple paths in thebus. Thus, the illustrative embodiments provide a method and apparatusfor controlling flight control surfaces on an aircraft.

A flight control system may contain a data bus system, an actuatorcontrol, and individual mixers in communication with each actuator. Thedata bus system is located in aircraft 100, and may be a part of aflight control computer and/or ailevatoron mixer 220.

The actuator control modules are connected to the data bus system. Anactuator control in the actuator may control positioning of a group offlight control surfaces on aircraft 100 using commands via the data bussystem that are directed to the actuator. Flight control programs and/orschedules may be connected to the data bus system. The flight controlprograms may generate and send the commands onto the bus system tocontrol the flight control surfaces on the aircraft. The commands for aflight control surface may be directed towards a group of actuatorcontrol programs on processors assigned to the actuators of the flightcontrol surfaces.

Flight control surfaces such as reduced horizontal stabilizer 132 and/orailevatoron 138 may be controlled by actuators which may be implementedin software, hardware, firmware or a combination thereof. When softwareis used, the operations performed by actuator controllers may beimplemented in program code configured to run on hardware, such as aprocessor unit. When firmware is used, the operations performed byactuator control programs and/or flight control programs, which may beimplemented in program code and data and stored in persistent memory torun on a processor unit. When hardware is employed, the hardware mayinclude circuits that operate to perform the operations in actuatorcontrol programs and flight control programs.

In the illustrative examples, without limitation the hardware for theprocessor units may take the form of a circuit system, an integratedcircuit, an application specific integrated circuit (ASIC), aprogrammable logic device, or some other suitable type of hardwareconfigured to perform a number of operations. With a programmable logicdevice, the device may be configured to perform the number ofoperations. The device may be reconfigured at a later time or may bepermanently configured to perform the number of operations. Examples ofprogrammable logic devices that may be used for processor units include,for example, a programmable logic array, a programmable array logic, afield programmable logic array, a field programmable gate array, andother suitable hardware devices. Additionally, the processes may beimplemented in organic components integrated with inorganic componentsand may be comprised entirely of organic components excluding a humanbeing. For example, the processes may be implemented as circuits inorganic semiconductors.

As depicted, each of flight control programs may part of processor unitsthat are dissimilar to each other and each of actuator control programmay include processor units that are dissimilar to each other. Forexample, one processor unit in the module may be implemented using acomputer microprocessor while the other processor in the module may beimplemented using a digital signal processor. As another example, twocomputer microprocessors may be used having different processorarchitectures.

The illustrations are not meant to imply physical or architecturallimitations to the manner in which an illustrative embodiment may beimplemented. Other components in addition to or in place of the onesillustrated may be used. Some components may be unnecessary. Also, theblocks are presented to illustrate some functional components. One ormore of these blocks may be combined, divided, or combined and dividedinto different blocks when implemented in an illustrative embodiment.

For example, a portion of a flight control system may use conventionalcontrols in addition to a fly-by-wire system. Further, a flight controlcomputer may control other types of devices other than flight controlsurfaces discussed in the figure. For example, the flight controlcomputer also may control the engines on aircraft 100.

As yet another example, a network may be used in addition to or in placeof data bus system to provide communications between actuator controlprograms and/or flight control programs. Further, some number of flightcontrol sub-programs may be used as supplements to the systems andprograms described herein for some illustrative examples. The operationsillustrated in FIGS. 6-9 may be implemented in aircraft 100 described inFIG. 1.

Operations shown receiving a signal and/or a command from an inputdevice may be from a of flight control in a cockpit of aircraft 100.These signals may be analog signals, digital signals, some combinationthereof, or signals transmitted mechanically via a cable, pully,linkage, or similar device. These signals may be generated from flightcontrols such as a flight stick, rudder pedals, a throttle, or someother suitable type of control. Flight controls may be controls operatedby a pilot, and/or by another operator and/or system within or datalinked from outside aircraft 100. In other illustrative examples, theflight controls may be devices sending sensor data or other informationneeded in the flight control modules to provide automatic adjustments tothe flight of aircraft without input from the pilots. Flight controldevices may send the commands onto a data bus system that may be withinor in communication with ailevatoron mixer 220 for flight controloperations.

Thus, embodiments above describe at least a process and machine forreducing a drag component of a horizontal stabilizer on an aircraft.More specifically, a process for reducing a size of a horizontalstabilizer for a particular model of an aircraft, may include augmentinga nose-up moment, for the particular model of the aircraft, provided bya reduced horizontal stabilizer for the particular aircraft model. Thisprocess may reduce the weight of the horizontal stabilizer for theparticular aircraft model via the reduced horizontal stabilizer for theparticular aircraft model. The process may include augmenting thenose-up moment via a deflection, away from a takeoff surface, of anailevatoron on the aircraft, which may involve deflecting symmetrically,away from a takeoff surface, each ailevatoron on the aircraft.

Additionally, the process disclosed in the embodiments may include aforce generated by the deflection the ailevatoron being located aft ofan axis of contact of the main gear of the aircraft on a takeoffsurface, and deflecting symmetrically, away from a takeoff surface, eachailevatoron on the aircraft based upon a degree of deflection of anelevator on the reduced horizontal stabilizer. Deflecting symmetrically,away from a takeoff surface, each ailevatoron on the aircraft may bebased upon a degree of deflection of an elevator on the reducedhorizontal stabilizer according to a determination in ailevatoron mixer,and/or upon a degree of deflection of an elevator on the reducedhorizontal stabilizer according to a determination in an ailevatoronmixer considering at least: a weight-on-wheels signal, an airspeed ofthe aircraft.

The process described by the embodiments above may also include basingthe size of the reduced horizontal stabilizer, for the particularaircraft model upon a rotation speed based upon an allowable combinationof gross takeoff weight and a forward center of gravity that results ina maximum nose-down moment about a point where main gear of theparticular aircraft model contact a takeoff surface. Reducing the sizeof the horizontal stabilizer for a particular aircraft model may occurby replacing the horizontal stabilizer, after manufacturing theparticular aircraft model using the horizontal stabilizer, with thereduced horizontal stabilizer, during at least one of: a modification, areconfiguration, a refurbishment, other maintenance, and otherservicing, of the particular aircraft model.

Further still, the above embodiments describe a process for increasingthe fuel efficiency for a particular aircraft model via reducing a sizeof a horizontal stabilizer for a particular aircraft model.

Increasing the fuel efficiency for a particular aircraft model mayinclude augmenting a nose-up moment, for the particular aircraft model,provided by a reduced horizontal stabilizer, and reducing the weight ofthe horizontal stabilizer for the particular aircraft model via thereduced horizontal stabilizer. Augmenting a nose-up moment provided bythe reduced horizontal stabilizer may occur via deflecting, away from atakeoff surface, a trailing edge of an ailevatoron on the aircraft. Theprocess may include deflecting symmetrically, away from a takeoffsurface, each ailevatoron on the particular aircraft model, andaugmenting the nose-up moment via generating a force, via theailevatoron being located aft of an axis of contact of the main gear ofthe particular aircraft model on a takeoff surface. Deflectingsymmetrically, away from a runway, each ailevatoron on the aircraft maybe based upon a degree of deflection, of an ailevatoron on the reducedhorizontal stabilizer, according to a determination by an ailevatoronmixer. Deflecting symmetrically, away from a runway, each ailevatoron onthe particular aircraft model may be based upon a degree of deflectionof an elevator on the reduced horizontal stabilizer according to adetermination by an ailevatoron mixer considering at least: aweight-on-wheels signal, and an airspeed of the particular aircraftmodel.

Additionally, the process disclosed in the embodiments may includebasing the size of the reduced horizontal stabilizer, to overcome amaximum nose-down moment for the particular aircraft model, resultingfrom a particular combination of gross takeoff weight and a maximumallowable forward center of gravity, upon a rotation speed for theparticular aircraft model. With this process, reducing the size of thehorizontal stabilizer for a particular aircraft model may occur byreplacing the horizontal stabilizer, after manufacturing the particularaircraft model using the horizontal stabilizer, with the reducedhorizontal stabilizer during at least one of: a modification, areconfiguration, a refurbishment, other maintenance, and otherservicing, of the particular aircraft model.

Embodiments above also describe a machine configured to reduce a size ofa horizontal stabilizer for a particular aircraft model while sustaininga nose-up moment, required for takeoff for the particular aircraft modelwith a maximum allowed nose-down moment about an axis where main gear ofthe particular aircraft model contact me takeoff surface, such that themachine comprises an ailevatoron mixer configured to symmetricallydeflect each ailevatoron, located aft of the axis, away from the takeoffsurface. The machine disclosed in the embodiments may also include theailevatoron mixer containing program code fixed in a non-transitorymedium configured to blend symmetric deflection, of each ailevatoronaway from the takeoff surface, based upon at least: deflection of anelevator on the reduced horizontal stabilizer, an airspeed, and aweight-on-wheels signal, of the particular aircraft model.

In addition to the embodiments shown above, additional embodiments forthe aileron mixer may direct symmetric deflection of each ailevatoron138 in a direction other than toward a surface 102 during a takeoff.Embodiments shown below recognize and take into consideration that asize requirement for horizontal stabilizer 116 may also be driven byrequirements to produce a nose-down pitch moment for an aircraft as itapproaches a high-angle-of-attack and/or stall condition. In otherwords, it may be possible to further reduce the size of reducedhorizontal stabilizer 132 when additional technical effects are added toimprove performance characteristics for a particular model of anaircraft in flight. As a non-limiting example, when an aircraftapproaches a high angle of attack or a stall in flight, the aircraftmust be able to provide a nose-down pitching moment that reduces theangle of attack or prevents the stall of the aircraft.

Embodiments shown below recognize that pitch control moments, measuredabout an aircraft's center of gravity, that are provided by a forcegenerated by a horizontal stabilizer, which is mounted aft a center ofgravity of an aircraft, will be reduced, for any given resultant forcegenerated by the horizontal stabilizer, as the aircraft center ofgravity moves back toward the horizontal stabilizer, reducing a staticstability for the aircraft, and shortening a moment arm from the centerof gravity to the horizontal stabilizer. Further reduction la nose-downpitch authority from a horizontal stabilizer may result when the centerof gravity is moved aft to or behind the aircraft aerodynamic center.Still further reduction in nose-down pitch authority from a horizontalstabilizer may result with the aircraft at high angles of attack, if afly-by-wire flight control system is added that can manage control foran aircraft with a relaxed static stability for the aircraft.

Embodiments described herein, recognize and take into account that anaircraft with relaxed static stability may have the center of gravitylocated at the aerodynamic center of the aircraft for neutral staticstability, or even have the center of gravity located aft of theaerodynamic center of the aircraft, such that the aircraft may bestatically unstable. Thus, depending upon a desired handlingcharacteristic for the aircraft, and a particular combination of flightcontrol surfaces for the aircraft, some required value of nose-downpitch moment for the aircraft, an aft limit is established based upon onan allowable range of locations for the aircraft center of gravity.

An airplane must be designed to operate within a range of a forward andan aft center of gravity location limit that accommodates movement ofthe center of gravity at least due to fuel burn and payload changes,without limitation. Thus, to minimize the average tail load over thise.g. range and thereby, on average, minimize trim drag, it would beadvantageous if the static stability could be relaxed beyond the neutralpoint.

The most fuel-efficient location of the center of gravity of theparticular aircraft (particularly for swept winged commercial transportaircraft models) for in-flight fuel efficiency, is at the aerodynamiccenter of the particular aircraft. Hence, dependent upon particularaircraft components and configurations and desired handlingcharacteristics of the particular aircraft, it is desirable for rangefor the allowable center of gravity off the particular aircraft toinclude a location at the aerodynamic center of the particular aircraft,such that as much of the flight as possible can be flown with the centerof gravity at or near the aerodynamic center of the particular aircraft.

Looking now to FIG. 10, FIG. 10 shows a plot of an angle of attackversus a maximum nose-down pitching moment coefficient (Cm) for anexample configuration of an example aircraft in-flight, depicted inaccordance with an advantageous embodiment. More specifically, FIG. 10shows chart 1000 with a vertical axis 1002 that represents an angle ofattack of a wing for a particular model of an aircraft in flight with acenter of gravity set at a point that gives the example aircraft relaxedstatic stability (RSS). Increasing angle of attack is shown movingupward on the vertical axis. Horizontal axis 1004 represents a maximumnose-down pitching moment coefficient (Cm) for the example aircraft. Byconvention, a pitching moment in a nose-down direction has a negativevalue. On chart 1000, Cm becomes more negative moving right on thehorizontal axis to indicate increasing magnitudes of a nose-downpitching moment for the aircraft.

For any particular model of aircraft, laws of aerodynamics and/orgovernment regulations may dictate certain in-flight performancerequirements for various configurations of the aircraft. Withoutlimitation, regulations may be aircraft airworthiness certificationregulations, such as without limitation, 14 USC § 25.

On chart 1000, line 1006 represents, for an example aircraft in anin-flight configuration with a center of gravity located at a maximumaft position in an allowable center of gravity range for the exampleaircraft, a pitch authority line that represents a nose-up pitch forceacting upon the example aircraft. When engine thrust produces a nose-uppitch moment, the nose-up pitch moment would normally plot to the leftof vertical axis 1002 in the positive Cm zone. For ease of visuallycomparing the nose-up pitch, due to thrust that must be overcome install recovery, to available nose-down pitch capability for the exampleaircraft, line 1006 represents a numerical negative (x −1) of linerepresenting a nose-up pitching moment due to thrust on the exampleaircraft.

An exact shape and location of line 1006 would thus be influenced by anamount of thrust each engine produces. Hence, line 1006 may berepresentative of a particular thrust condition for the exampleaircraft. Accordingly, further shifts in location on chart 1000 or shapechanges for line 1006 may result from a thrust limiting system on theaircraft. A shifting of line 1006 to the left may be especiallynoticeable for an aircraft at light weight with a far aft center ofgravity, where specific excess power [p_(s)≈V·(cos(α)·T−drag)/W)]provides more than enough climb capability, where V equals airspeed, Tequals thrust, and W equals a weight of the aircraft.

On vertical axis 1002, alpha-limit (α_(limit)) 1008 indicates an angleof attack for the wing of the example aircraft where an undesiredperformance of the aircraft may begin, such as without limitationairflow separation, and represents an angle of attack that the aircraftdesigner/operator prefers not be exceeded in flight. The angle of attackthat the aircraft designer/operator prefers not be exceeded in flight,α_(limit) 1008 may be maximum angle of attack required to meetregulatory requirements, such as without limitation. Federal AviationRegulation (FAR) Part 25 maneuver requirements. Alpha-stall (α_(stall))1010 represents an angle of attack at which the aircraft is consideredto be in a stall condition, as described at least in FAR Part 25 §§203-207, and § 103. In order for the example aircraft to avoid enteringan unrecoverable stall, the example aircraft must maintain a nose-downpitch authority that is to the right of line 1006.

Curve 1012 represents a nose-down pitch authority provided by ahorizontal stabilizer on the example aircraft. If at any point, curve1012 were to touch line 1006, the example aircraft would have zeronose-down recovery capability, and if curve 1012 were shift further tothe left, such that a portion of curve 1012 were to the left of line1006, for the angles of attack where curve 1012 is to the left of line1006 it would indicate a non-recoverable hung stall condition. In otherwords, the horizontal stabilizer could not provide enough nose-downpitch to move the aircraft out of the stalled angle of attack. Hence,there may be a minimum differential that is desired between lines 1006and 1012. This minimum differential may be driven by overall aircraftdesign considerations, and operator preference, or by regulation.Preferences and regulations for the minimum differential may varydependent, without limitation, upon a category, a phase of flight, aconfiguration, and/or a technical effect provided by and/or anoperational status of a component on the aircraft that affectsperformance capability of a flight control system, of the aircraft.

The leftward slope of curve 1012 changes above α_(limit) 1008 for theexample aircraft of chart 1000 because, as is common to many transportcategory aircraft with a tail mounted horizontal stabilizer, aboveα_(limit) 1008 airflow disturbances and/or separation occurring on thewing of the example aircraft interfere with desired airflow over thehorizontal stabilizer, and thus reduce the aerodynamic force which thehorizontal stabilizer can produce, thereby reducing a magnitude of theCm produced by the horizontal stabilizer. Thus, α_(limit) 1008 indicatesan angle of attack, above which operation of an aircraft, such aswithout limitation aircraft 100, may net be desired during normaloperation. Operation above α_(limit) 1008 may not be desired at leastbecause of excessive buffet from airflow separating off of the wings ofaircraft 100, and the closeness to the stall angle of attack.Maneuvering an example aircraft represented on chart 1000 at angles ofattack greater than α_(limit) 1008 may not be required. Maneuvering anexample aircraft represented on chart 1000 at angles of attack greaterthan α_(limit) 1008 may not be required for certification of theparticular model of the example aircraft.

The location, shape, and slopes of curve 1012 are influenced at least bya size, a location, and a shape of the horizontal stabilizer thatproduces the curve plots. If the size of the horizontal stabilizerproducing curve 1012 were reduced in size, the smaller horizontalstabilizer would produce the curve 1016.

In chart 1000, distance 1014 represents a desired and/or requiredminimum differential between a Cm value along line 1006 and a Cm valuealong curve 1012, at a given angle of attack. Distance 1014 may providea margin of desired and/or required minimum nose-down recoverycapability for the aircraft in the conditions represented by chart 1000.Thus, given distance 1014 as a minimum, curve 1012 cannot be moved anyfurther to the left on chart 1000 if the desired/required aircraftnose-down recovery capability are to be sustained.

One of skill in the art can see that on chart 1000, curve 1016 does notmaintain distance 1014 from line 1006 at (α_(stall)) 1010. Hence, onecan see that when reducing the size of the horizontal stabilizer thatcan produce curve 1012 to a smaller size that produces curve 1016 wouldnot be acceptable.

However, ailevatoron mixer 220 provides a machine that provides thetechnical effect of augmentation to the Cm produced by the horizontalstabilizer producing curve 1016 up to the Cm performance represented bycurve 1012. In other words, for an aircraft in flight, ailevatoron mixer220 can augment nose-down pitch authority for aircraft 100 in a mannersimilar to—but in an opposite (nose-down) direction—to the ailevatoronmoment 142 used to augment reduced moment 136 at aircraft rotation ontakeoff.

For either curve 1016 or curve 1012 a horizontal distance from line 1006to curve 1016 or curve 1012 is an indication of excessive capability forsmaller horizontal stabilizer 1112 or horizontal stabilizer 116respectively, and thus excessive size, weight, and/or drag of smallerhorizontal stabilizer 1112 or horizontal stabilizer 116 respectivelyabove the size, weight, and/or drag required by line 1006 for theexample aircraft. While chart 1000 is an example for an example of someparticular aircraft model with a set center of gravity in a particularconfiguration with a particular thrust level, the concepts discussedhere are applicable for similar charts for the example aircraft indifferent configurations at different thrusts, and may apply likewisefor other aircraft models with particular set conditions. While aprecise shape location or slope of line 1006 may change, itssignificance and relationship to lines 1012 and 1016 remainrepresentative of a nose-up pitch that forms a limit which must beovercome with pitch command authority for the particular aircraft modelto be certified as having acceptable stall recovery characteristics.

With reference now to FIG. 11, FIG. 11 shows an aircraft in flight witha wing at a stall angle off attack, depicted in accordance with anadvantageous embodiment. More specifically, FIG. 11 shows aircraft 100in flight with wing 126 at α_(stall) 1010 in airflow 1102. Ailevatoronmixer 220 receives inputs of pitch command 232 and configuration ofdeflection 234 of elevator 226. Acceleration 242 may include not only achange in airspeed of aircraft 100, but changes in groundspeed, andnormal acceleration loads, or “g′s” experienced by aircraft 100. Incontrast to scheduler 250 output at rotation for symmetric deflection252 of ailevatoron(s) 138 in a direction that produces a nose-upailevatoron moment 142 for aircraft 100, as discussed above for FIGS. 1and 2, when aircraft 100 is in-flight, and weight on wheels signal 222received by filter 236 within ailevatoron mixer 220 indicates no weighton wheels, and pitch command 232 for deflection 234 of elevator 226 is afull nose-down pitch command 232 for aircraft 100, and/or deflection 234of elevator 226 is at a full nose-down position, then output ofscheduler 250 directs a full symmetric deflection 252 of ailevatoron(s)138 in a direction that produces ailevatoron moment 1104 in a nose-downdirection about center of gravity 1106 for aircraft 100. Center ofgravity 1106 may be located at a full aft position within range 112 ofallowable locations for center of gravity 1106 for aircraft 100.

In contrast to rotation of aircraft 100 at takeoff from surface 102,where a forward most center of gravity 110 is likely to dictate aminimum size allowable for a horizontal stabilizer, while aircraft 100is in flight, an aft-most location in range 112 for center of gravity1106 is likely to dictate a minimum size allowable for a horizontalstabilizer. Thus, for an aircraft without a fly-by-wire flight controlsystem, a mechanical linkage and/or a mechanical feedback system can beused to inform ailevatoron mixer 220 that a full nose-down pitch command232 exists and/or that deflection 234 of elevator 226 is at a fullnose-down position, and thus trigger the symmetrical deflection 252 ofeach ailevatoron 138 in a full nose-down, for aircraft 100, direction,full trailing edge down (TED) symmetric deflection 252 for eachailevatoron 138 to generate force 1108 that generates ailevatoron moment1104.

For an aircraft lacking fly-by-wire stability augmentation, the aircraftwould most likely be designed without relaxed static stability, and thuscurve 1012 would more likely have a shape and slope similar to thatshown by curve 1306 in FIG. 13. In other words, ailevatoron mixer 220may generate ailevatoron moment 1104 to augment smaller moment 1110produced by smaller horizontal stabilizer 1112 to a magnitude equal toor greater than stab moment 1114 (produced by larger sized horizontalstabilizer 116) which is greater than nose-up moment 1116. Nose-upmoment 1116 may be associated with line 1006 presented in FIG. 10.Smaller horizontal stabilizer 1112 may be identical to reducedhorizontal stabilizer 132, or may be a different size and/or shape fromreduced horizontal stabilizer 132.

The size and shape of reduced horizontal stabilizer 132 derives fromdesign considerations for the rotation requirements for takeoff, whereasa shape and a size for smaller stabilizer 132 are driven by in-flightperformance considerations. One or the other may dominate in designconsiderations based upon desired characteristics for aircraft 100takeoff and in-flight performance, and/or some combination/compromisebetween the requirements for the two phases of operation for aircraft100 may be selected.

For a particular aircraft model, a wing configuration and location and amain gear location that accommodate a required range of travel for thecenter of gravity throughout a flight, such that neither the forward oraft center of gravity limits dominate a design of a size for thehorizontal stabilizer that may be desired and/or selected. In otherwords, in an optimal design, a single sized horizontal stabilizer willmeet both nose-up and nose-down moment requirements for the aircraftloaded at forward-most and aft-most allowable locations for the centerof gravity, respectively. Applying ailevatoron mixer 220 directedsymmetrical deflection 252 of each ailevatoron 138 in a full nose-down,for aircraft 100, direction, in conjunction with fly-by-wire flightcontrols and a zoom climb prevention system (ZCPS) as described inpatent application Ser. No. 14/921,841 titled Zoom Climb PreventionSystem for Enhancing Performance, filed Oct. 23, 2015 by the Applicant,allows for further advantages from the technical effects provided by thecombination of the ZCPS, fly-by-wire flight controls, and ailevatoronmixer 220.

Accordingly, application Ser. No. 14/921,841 titled Zoom ClimbPrevention System for Enhancing Performance, filed Oct. 23, 2015 by theApplicant is incorporated herein in its entirety. Thus, when a ZCPS isactive on aircraft 100, such that aircraft 100 would not exceed an angleof attack represented by α_(limit) 1008, then it may be possible to takerelief from the required minimum differential between a Cm value alongline 1006 and a Cm value along curve 1012, α_(stall) 1010, such that itmay be acceptable for curve 1012 to be tangent to or slightly right of(less than distance 1014) line 1006.

Embodiments herein recognize and take into account that regardless ofwhether aircraft 100 has a fly-by-wire flight control system or not, inflight at angles of attack above α_(limit) 1008 some loss of amount offorce producible by an aileron on an aircraft may occur due toseparation of airflow from wing 126, dependent upon a particularlocation and/or size of the aileron on wing 126 and the particularairflow response of the particular design for wing 126. The loss ofamount of force producible by an aileron on an aircraft that may occurin flight at angles of attack above α_(limit) 1008 due to separation ofairflow from wing 126 may be a significant amount.

Hence, to generate desired forces and/or to optimize effectiveness ofailevatoron mixer 220, in some embodiments, previously designed aileronsmay be replaced by ailevatoron(s) 138 that may have at least one off: adifferent location along wing 126, a different size, a differentallowable degree of trailing-edge deflection up and/or down, anadditional number of ailevatoron(s) 138 added onto wing 126, and/or anycombinations thereof. In other embodiments, and particularly whenaircraft 100 includes a fly-by-wire flight control system thatincorporates a zoom climb prevention system that prevents angle ofattack for aircraft 100 in flight from going above α_(limit) 1008,ailevatoron mixer 220 may control previously existing ailerons asailevatoron(s) 138.

With reference now to FIG. 12A, FIG. 12A shows a plot of an angle ofattack versus a maximum nose-down pitching moment coefficient (Cm) forthe example configuration of the example aircraft in-flight, depicted inaccordance with an advantageous embodiment. More specifically. FIG. 12Ashows the same vertical, angle of attack, and horizontal, maximumnose-down pitching moment (Cm), axes, line 1006, curve 1012, anddistance 1014 as presented in FIG. 10. Curve 1202 represents a shift incurve 1012 that may be allowable for an aircraft utilizing the ZCPS.Curve 1012 may represent Cm plots for a horizontal stabilizer 116 asshown in FIG. 11.

When equipped with the ZCPS, aircraft 100 should never exceed α_(limit)1008 in flight. With aircraft 100 designed to never exceed α_(limit)1008 and required distance 1014 being present at α_(limit) 1008 betweenline 1006 and curve 1202, FIG. 12A indicates that a smaller horizontalstabilizer 1112 (as shown in FIG. 11) may be used on aircraft 100 inplace of horizontal stabilizer 116 represented by curve 1012. Atechnical effect of this shift of curve 1012 to curve 1202 is that forall angles of attack, a distance between line 1006 and curve 1202 isless than a distance between line 1006 and curve 1012.

Hence, a technical effect of using ZCPS may be to allow use of smallerhorizontal stabilizer 1112 throughout a flight envelope for aircraft 100and thus reduce a weight, drag, and fuel consumption throughout a flightenvelope for aircraft 100. Alternatively, if regulatory and/or otherconsiderations do not allow using a horizontal stabilizer that has a Cmcurve as depicted by curve 1202, regulatory or ether considerations mayallow a horizontal stabilizer that has a Cm curve as depicted by curve1202 when the aircraft includes ailevatoron mixer 220 and the capabilityto augment a Cm for the horizontal stabilizer of a size that has a Cmcurve as depicted by curve 1202.

Still further, it may be possible that regulatory and/or other designconsiderations would allow for near zero or just slightly positivenose-down pitch capability at α_(stall) 1010. With reference now to FIG.12B, FIG. 12B shows a plot of an angle of attack versus a maximumnose-down pitching moment coefficient (Cm) for the example configurationof the example aircraft of FIG. 12A in-flight, depicted in accordancewith an advantageous embodiment.

More specifically, in FIG. 12B, chart 1201 shows that if, at leastbecause of reduced risk that an aircraft may exceed an α_(limit) 1008,regulatory and/or other design considerations would allow curve 1202 tobe shifted even further left, as indicated by curve 1204, such that whenaircraft 100 maximum nose-down pitch authority is augmented byailevatoron mixer 220 and ailevatoron(s) 138, Cm performance foraircraft 100 may be modified as indicated by curve 1206, then a size forsmaller horizontal stabilizer 1112 may be reduced even further thanindicated in FIG. 12A by curve 1202. Thus, a technical effect ofcombining the ailevatoron mixer 220 in communication with a fly-by-wireflight control system utilizing a zoom climb prevention system (ZCPS)will be an even greater reduction to a size of horizontal stabilizer116, and thus an even greater reduction of weight and/or drag, and/orfuel consumption for aircraft 100, as compared to aircraft 100 utilizingonly ailevatoron mixer 220 and ailevatoron(s) 138 as shown in FIG. 10.

One of skill in the art will recognize that depending upon particularstall characteristics for a particular wing, that a size and a number ofailevatoron 138 located on each wing, and particular locations for eachailevatoron 138 on a trailing edge of each wing, may affect theparticular shape of curve 1206 shown in FIG. 12B, and/or curve 1012 inFIGS. 10 and 12A. Thus, in conjunction with features and programmedfunctions of ailevatoron mixer 220, various designs for a size, a shape,and a location for each ailevatoron 138 may be tailored to provide adesired shape for curves 1012 and 1206 in FIGS. 10, 12A and 12B.

One of skill in the art recognizes that for all angles of attack at orbelow α_(limit) 1008, that curve 1204 maintains a distance from line1006 that approaches or exceeds distance 1014 currently required foraircraft without an ailevatoron mixer 220 or a ZCPS. Hence, if α_(limit)1008 will not be exceeded, and nose-down pitch authority is maintainedfor aircraft 100 at all expected angles of attack, then when horizontalstabilizer 116 is reduced to smaller horizontal stabilizer 1112, thetechnical effect is a reduction in weight and drag for aircraft 100 anda resultant increase in performance to include a fuel efficiency.

One of skill in the art will also recognize and take into considerationthat FIGS. 10, 12A and 12B are representative for performance of anexample aircraft in an example configuration, with a center of gravityset such that the example aircraft has a relaxed static stability.However, the ability to augment Cm provided by ailevatoron mixer 220 maybe used instead of, or in combination with, reducing a size of ahorizontal stabilizer for the example aircraft, to utilize an evenfurther aft center of gravity.

While charts 1000, 1200, and 1201 present examples for someexample/notional aircraft model with a set center of gravity in aparticular configuration with a particular thrust level, the conceptsdiscussed here are applicable for comparable charts for the exampleaircraft in different configurations, with different horizontalstabilizer trim settings, or at different thrusts, and may applylikewise for other aircraft models with particular set conditions. Whilea precise shape location or slope of line 1006 may change, itssignificance and relationship to lines 1012 and 1016 remainrepresentative of a nose-up pitch that forms a limit which must beovercome with pitch command authority for the particular aircraft modelto be certified as dynamically stable.

with reference now to FIG. 13, FIG. 13 shows a plot of an angle ofattack versus a maximum nose-down pitching moment coefficient (Cm) foran example aircraft in-flight with a center of gravity located atvarious positions, depicted in accordance with an advantageousembodiment. More specifically, FIG. 13 presents chart 1300 with the samevertical axis 1002 and horizontal axis 1004 presented in FIGS. 10, 12A,and 12B. Line 1006 also remains, and shows, for an example aircraft inan in-flight configuration with a center of gravity located at a maximumaft position in an allowable center of gravity range for the exampleaircraft, a pitch authority line for a pitch force acting upon anaircraft. Line 1006 represents a negative of a nose-up pitching momentdue to thrust on the example aircraft.

In addition to shifting, curve 1012 to the left on chart 1300 viareducing a size of a horizontal stabilizer for the example aircraft, oreffectively shifting curve 1012 to the right via the added pitchauthority provided by ailevatoron mixer 220 and symmetrical deflection252 of ailevatoron(s) 138, as shown above in FIGS. 10, 12A, and 12B,curve 1012 may shift by changing its slope to tilt further leftward,which would indicate further relaxing the static stability for theexample aircraft represented by curve 1012.

As mentioned above, changes to the aircraft may change the shape of thecurves on the charts. Changing a location of the center of gravity ofthe example aircraft relative to the mean aerodynamic chord of theexample aircraft, changes the slope of curve 1012, and provides a slightshift in where it crosses axis 1004 due to the reduced horizontalstabilizer moment arm. By allowing a further aft center of gravity thancurve 1012 may slope further to the left and shift slightly left on axis1004. As a non-limiting example, the shape of curve 1012 may berepresentative for a notional aircraft with a tail mounted horizontalstabilizer, and an aft center of gravity of the notional aircraftlocated at 40% mean aerodynamic chord providing a relaxed staticstability, while curve 1302 may be representative of the same notionalaircraft with a same horizontal stabilizer, but with the center offgravity off the notional aircraft located further aft, at 60% meanaerodynamic chord. Hence, embodiments of a machine and/or a processdescribed above also can provide for moving the center off gravityfurther aft, and produce the technical effect of reducing a dragcomponent off a tail mounted horizontal stabilizer during flight, andthus increasing a fuel efficiency of the aircraft.

Without relaxing stable stability, a center of gravity may be located at20% mean aerodynamic chord and represented by curve 1306. As mentionedabove, the slope and shape of curve 1302 may be more prevalent on anaircraft devoid of a fly-by-wire flight control system.

In chart 1300, curve 1304 indicates enhanced nose-down pitch authorityprovided by ailevatoron mixer 220 and ailevatoron 138(s). Hence, evenwith a center of gravity located back at 60% mean aerodynamic chord, atα_(stall) 1010, and unaugmented performance shown by curve 1302,ailevatoron mixer 220 and ailevatoron 138(s) augmented pitch nose-downauthority indicated by curve 1304 remains to the right of line 1006 andprovides acceptable stall recovery capability for the aircraft byavoiding a non-recoverable hung stall condition indicated by maximumnose-down pitching moment coefficient values to the left of line 1006.

Thus, even without changing a size of a horizontal stabilizer, utilizingailevatoron mixer 220 to augment nose-down pitch authority of anaircraft 100 produces the technical effect of expanding the operatingenvelope for the aircraft. By allowing a further aft center of gravity,a larger range for allowable center of gravity location, and thus alarger variation in payload distributions may be carried by aircraft 100with ailevatoron mixer 220 as compared to aircraft 100 operating withoutailevatoron mixer 220, and/or ZCPS. An allowable center of gravity foran aircraft utilizing ailevatoron mixer 220 may be further aft relativeto mean aerodynamic chord than for the same aircraft devoid ofailevatoron mixer 220. Without limitation, an aircraft model utilizingailevatoron mixer 220 may gain the benefit from technical effectsresulting from operating with a center of gravity location up to 5%further aft relative to a mean aerodynamic chord of the aircraft modeldevoid of ailevatoron mixer 220.

For a nominal aircraft that combines use of a ZCPS with ailevatoronmixer 220, based upon stall margin requirements, a technical effect maybe the ability to operate with a center of gravity for the nominalaircraft moved as far aft as 65% of mean aerodynamic chord, dependentupon regulatory approval and other design requirements for the nominalaircraft. For a nominal aircraft that combines use of a ZCPS withailevatoron mixer 220 may gain the benefit from technical effectsresulting from operating with a center of gravity location 5% to 10%further aft relative to a mean aerodynamic chord of the aircraft modeldevoid of ailevatoron mixer 220.

Hence, above has been described at least a process for improving aperformance of a particular model of an aircraft via reducing a size ofa horizontal stabilizer on the particular model of the aircraft, thusforming a smaller horizontal stabilizer, and sustaining a required stallrecovery nose-down pitching moment capability of the aircraft. Thus theembodiments above describe at least a machine that includes anailevatoron mixer configured to symmetrically deflect an ailevatoron oneach wing of an aircraft and thereby improve a performance of theaircraft compared to the aircraft configured without the ailevatoronmixer. The machine may also include the aircraft having an improvementin performance that includes an increase in a fuel efficiency of theaircraft as compared to the aircraft operating without the ailevatoronmixer.

The aircraft may also include an improvement in the performance ascompared to the aircraft operating without the ailevatoron mixer, via areduction in a size and a weight of a horizontal stabilizer. Theaircraft may also include an improvement in the performance as comparedto the aircraft operating without the ailevatoron mixer, wherein theimprovement includes at least one of: an extension of a range allowedfor a location of a center of gravity of the aircraft, and a reductionin a size and a weight of a horizontal stabilizer combined therewith.Additionally, the machine may include the ailevatoron mixer configuredto communicate with the ZCPS.

Hence, the embodiments above describe the machine as usable at least ina process for improving a performance of a particular model of anaircraft. Looking now to FIG. 14, FIG. 14 presents a flow chart for aprocess for improving a performance of a particular model of an aircraftvia reducing a size of a horizontal stabilizer on the particular modelof the aircraft to form a smaller horizontal stabilizer, and sustaininga required stall recovery nose-down pitching moment capability of theaircraft, in accordance with an illustrative embodiment.

More specifically, process 1400 may begin by reducing a size of ahorizontal stabilizer on the particular model of the aircraft to form asmaller horizontal stabilizer, and sustaining a required stall recoverynose-down pitching moment capability off the aircraft (operation 1402).The required stall recovery nose-down pitching moment capability of theaircraft may be for the aircraft comprising an aft-most allowable centeroff gravity loading. When using an ailevatoron mixer, a size of ahorizontal stabilizer may be reduced compared to the size of ahorizontal stabilizer required for the same aircraft model withoutailevatoron mixer 220. Without limitation, use of ailevatoron mixer 220and ailevatoron(s) 138 may provide the technical benefit of in-flightstall recovery performance enhancements that allow for reducing the sizeof the horizontal stabilizer by up to 3%. Without limitation, use ofailevatoron mixer 220 and ailevatoron(s) 138 in conjunction with a zoomclimb prevention system, may provide in-flight stall recoveryperformance enhancements that allow for reducing the size of thehorizontal stabilizer by 5-10%.

Process 1400 may also include improving the performance via reducing adrag component of the particular model of the aircraft (operation 1404).Process 1400 may also include improving the performance via increasing afuel efficiency of the aircraft (operation 1406). Process 1400 may alsoInclude reducing a weight of the horizontal stabilizer (operation 1408).

In conjunction with a zoom climb prevention system, process 1400 mayinclude improving a performance via a reduction in an approach airspeedfor the particular model of the aircraft (operation 1410). Process 1400may also include commanding a full aircraft nose-down deflection of anelevator on the smaller horizontal stabilizer (operation 1412).

Accordingly, process 1400 may include receiving in an ailevatoron mixer,the pitch command and direct symmetrically deflecting the ailevatoron oneach wing respectively to a full trailing edge downward position(operation 1414). Hence, the process may include sustaining a requiredstall recovery nose-down pitching moment capability of the aircraft viasymmetrically deflecting downward an ailevatoron located on each wing ofthe aircraft (operation 1416). Likewise, the process may also includesymmetrically deflecting downward the ailevatoron located on each wingof the aircraft responsive to a pitch command (operation 1418).

As a reminder, similar to the previous discussion re reduced horizontalstabilizer 132 discussed re FIG. 7 above, although smaller horizontalstabilizer 1112 may be designed into aircraft 100 before manufacturingbegins, reducing the size of horizontal stabilizer 116 for a particularaircraft model may occur after an original design phase for a particularaircraft model, or aircraft 100 in particular. Reducing the size ahorizontal stabilizer for a particular aircraft model, such as withoutlimitation, horizontal stabilizer 116 of aircraft 100, may occur byreplacing horizontal stabilizer 116 with smaller horizontal stabilizer132, after manufacturing the particular aircraft model with a particularsize horizontal stabilizer, during at least one of: a modification, areconfiguration, a refurbishment, other maintenance, and otherservicing, of the particular aircraft model (similar to operation 716 inFIG. 7).

Additionally, the embodiments above describe a machine usable at leastin a process for expanding an allowable range for a center of gravity ofa particular model of an aircraft and retaining a configuration of theparticular model of the aircraft. With reference now to FIG. 15, FIG. 15shows a flowchart for a process for expanding an allowable range for acenter of gravity of a particular model of an aircraft and retaining aconfiguration of the particular model of the aircraft.

More specifically, process 1500 may include expanding an allowable rangefor a center of gravity of a particular model of an aircraft and retaina configuration of the particular model of the aircraft (operation1502). Process 1500 may also include increasing a nose-down pinchauthority of the particular model of the aircraft (operation 1504).

Process 1500 may also include increasing a nose-down pitch authority ofthe particular model of the aircraft via an ailevatoron mixer providingsymmetrical deflection of an ailevatoron on each wing of the particularmodel of the aircraft (operation 1506). Hence process 1500 may includecommanding an increasing nose-down pitching moment for the aircraftresponsive to an elevator receiving a command for full nose-down pitch(operation 1508). Additionally, process 1500 may include the ailevatoronmixer controlling an ailevatoron symmetrical deflection in coordinationwith a zoom climb prevention system (operation 1510).

The flowcharts and block diagrams in the different depicted embodimentsillustrate the architecture, functionality, and operation of somepossible implementations of apparatuses and methods in an illustrativeembodiment. In this regard, each block in the flowcharts or blockdiagrams may represent at least one of a module, a segment, a function,or a portion of an operation or step. For example, one or more of theblocks may be implemented as program code, in hardware, or a combinationof the program cede and hardware. When implemented in hardware, thehardware may, for example, take the form of integrated circuits that aremanufactured or configured to perform one or more operations in theflowcharts or block diagrams. When implemented as a combination ofprogram code and hardware, the implementation may take the form offirmware.

In some alternative implementations of an illustrative embodiment, thefunction or functions noted in the blocks may occur out of the ordernoted in the figures. For example, in some cases, two blocks shown insuccession may be executed substantially concurrently, or the blocks maysometimes be performed in the reverse order, depending upon thefunctionality involved. Also, other blocks may be added in addition tothe illustrated blocks in a flowchart or block diagram.

Thus, the illustrative embodiments provide a method and apparatus formanaging commands for flight control surfaces. One or more illustrativeembodiments may use fly-by-wire systems for aircraft. The description ofthe different illustrative embodiments has been presented for purposesof illustration and description, and is not intended to be exhaustive orlimited to the embodiments in the form disclosed. Many modifications andvariations will be apparent to those of ordinary skill in the art.Further, different illustrative embodiments may provide differentfeatures as compared to other desirable embodiments. The embodiment orembodiments selected are chosen and described in order to best explainthe principles of the embodiments, the practical application, and toenable others of ordinary skill in the art to understand the disclosurefor various embodiments with various modifications as are suited to theparticular use contemplated.

What is claimed is:
 1. A process for improving a performance of aparticular model of an aircraft via: reducing a size of a horizontalstabilizer on the particular model of the aircraft to form a smallerhorizontal stabilizer; and sustaining a required snail recoverynose-down pitching moment capability of particular model of theaircraft.
 2. The process of claim 1, further comprising the requiredstall recovery nose-down pitching moment capability of the aircraftbeing for the particular model of the aircraft comprising an aft-mostallowable center of gravity loading.
 3. The process of claim 1, furthercomprising the performance comprising a reduction in a drag component ofthe particular model of the aircraft.
 4. The process of claim 1, furthercomprising the performance comprising an increase in a fuel efficiencyof the particular model of aircraft.
 5. The process of claim 4, furthercomprising a reduction of a weight of the horizontal stabilizer.
 6. Theprocess of claim 1, further comprising the particular model of theaircraft comprising a zoom climb prevention system and the performancecomprising a reduction in an approach airspeed for the particular modelof the aircraft.
 7. The process of claim 1, further comprisingsustaining a required stall recovery nose-down pitching momentcapability of the particular model of the aircraft comprisingsymmetrically deflecting downward an ailevatoron located on each wing ofthe aircraft.
 8. The process of claim 1, further comprisingsymmetrically deflecting downward an ailevatoron located on each wing ofthe particular model of the aircraft responsive to a pitch command. 9.The process of claim 8, further comprising the pitch command comprisinga command for a full aircraft nose-down deflection of an elevator on thesmaller horizontal stabilizer.
 10. The process of claim 8, furthercomprising an ailevatoron mixer receiving the pitch command anddirecting symmetrically deflecting the ailevatoron on each wingrespectively to a full trailing edge downward position.
 11. A processfor expanding an allowable range for a center of gravity of a particularmodel of an aircraft and retaining a configuration of the particularmodel of the aircraft.
 12. The process of claim 11, further comprisingincreasing a nose-down pitch authority of the particular model of theaircraft.
 13. The process of claim 11, further comprising increasing anose-down pitch authority of the particular model of the air-craft viaan ailevatoron mixer providing symmetrical deflection of an ailevatoronon each wing of the particular model of the aircraft.
 14. The process ofclaim 11, further comprising an ailevatoron mixer commanding anincreasing nose-down pitching moment for the particular model of theaircraft responsive to an elevator receiving a command for fullnose-down pitch.
 15. The process of claim 11, further comprising anailevatoron mixer controlling an ailevatoron symmetrical deflection incoordination with a zoom climb prevention system.
 16. A machine thatcomprises and ailevatoron mixer configured to symmetrically deflect anailevatoron on each wing of an aircraft and thereby improve aperformance of the aircraft compared to the aircraft configured withoutthe ailevatoron mixer.
 17. The machine of claim 16, further comprisingan improvement in performance comprising an increase in a fuelefficiency of the aircraft.
 18. The machine of claim 16, furthercomprising an improvement in the performance comprising a reduction in asize and a weight of a horizontal stabilizer.
 19. The machine of claim16, further comprising an improvement in the performance comprising atleast one of: an extension of a range allowed for a location of a centerof gravity of the aircraft, and a reduction in a sire and a weight of ahorizontal stabilizer combined therewith.
 20. The machine of claim 16,further comprising the ailevatoron mixer configured to communicate witha zoom climb prevention system.